2004 NASA Seal/Secondary Air System Workshop

2004 NASA Seal/Secondary Air SystemWorkshop

NASA/CP—2005-213655/VOL1

October 2005

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2004 NASA Seal/Secondary Air SystemWorkshop

NASA/CP—2005-213655/VOL1

October 2005

National Aeronautics andSpace Administration

Glenn Research Center

Proceedings of a conference held at Ohio Aerospace Institutesponsored by NASA Glenn Research CenterCleveland, OhioNovember 9–10, 2004

Available from

NASA Center for Aerospace Information7121 Standard DriveHanover, MD 21076

National Technical Information Service5285 Port Royal RoadSpringfield, VA 22100

Trade names or manufacturers’ names are used in this report foridentification only. This usage does not constitute an officialendorsement, either expressed or implied, by the National

Aeronautics and Space Administration.

Available electronically at http://gltrs.grc.nasa.gov

This work was sponsored by the Low Emissions AlternativePower Project of the Vehicle Systems Program at the

NASA Glenn Research Center.

Contents were reproduced from author-providedpresentation materials.

NASA/CP—2005-213655/VOL1 iii

Executive Summary Volume 1

The 2004 NASA Seal/Secondary Air System workshop covered the following topics: (i) Overview

of NASA’s new Exploration Initiative program aimed at exploring the Moon, Mars, and beyond; (ii) Overview of the NASA-sponsored Ultra-Efficient Engine Technology (UEET) program; (iii) Overview of NASA Glenn’s seal program aimed at developing advanced seals for NASA’s turbomachinery, space, and reentry vehicle needs; (iv) Reviews of NASA prime contractor and university advanced sealing concepts including tip clearance control, test results, experimental facilities, and numerical predictions; and (v) Reviews of material development programs relevant to advanced seals development.

The NASA UEET overview illustrated for the reader the importance of advanced technologies, including seals, in meeting future turbine engine system efficiency and emission goals. For example, the NASA UEET program goals include an 8- to 15-percent reduction in fuel burn, a 15-percent reduction in CO2, a 70-percent reduction in NOx, CO, and unburned hydrocarbons, and a 30-dB noise reduction relative to program baselines.

The workshop also covered several programs NASA is funding to develop technologies for the Exploration Initiative and advanced reusable space vehicle technologies. NASA plans on developing an advanced docking and berthing system that would permit any vehicle to dock to any on-orbit station or vehicle, as part of NASA’s new Exploration Initiative. Plans to develop the necessary mechanism and “androgynous” seal technologies were reviewed. Seal challenges posed by reusable re-entry space vehicles include high-temperature operation, resiliency at temperature to accommodate gap changes during operation, and durability to meet mission requirements.

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Contents Overview of NASA Glenn Seal Developments

Bruce M. Steinetz, Margaret P. Proctor, and Patrick H. Dunlap, Jr., NASA Glenn Research Center; Irebert Delgado, U.S. Army Research Laboratory; Joshua Finkbeiner, NASA Glenn Research Center; Jeffrey J. DeMange and Christopher C. Daniels, University of Toledo; and Scott B. Lattime, Ohio Aerospace Institute ....................................... 1

Overview of NASA’s Exploration Initiative Joseph Nainiger, NASA Glenn Research Center......................................................................... 31 Overview of the Ultra-Efficient Engine Technology and Quiet Aircraft Technology Projects Carol Ginty, NASA Glenn Research Center................................................................................ 47 High Misalignment Carbon Seals for the Fan Drive Gear System Technologies

Dennis Shaughnessy and Lou Dobek, United Technologies—Pratt & Whitney......................... 89 Leakage and Power Loss Test Results for Competing Turbine Engine Seals

Margaret P. Proctor, NASA Glenn Research Center; and Irebert R. Delgado, U.S. Army Research Laboratory.................................................................................................. 111

Test Rig for Evaluating Active Turbine Blade Tip Clearance Control Concepts: An Update

Scott B. Lattime, Ohio Aerospace Institute; Bruce M. Steinetz and Kevin Melcher, NASA Glenn Research Center; Jonathan A. Decastro, QSS; and Malcolm G. Robbie and Arthur H. Erker, Analex Corporation.................................................................................... 131

Wear Prediction of Strip Seals Through Conductance Norman Turnquist and Farshad Ghasripoor, GE Global Research; and Mark Kowalczyk and Bart Couture, GE Energy ...................................................................................................... 149

Parametrical Study of Hydrodynamic Seal Using a 2D Design Code and Comparing with a 3D CFD Model Xiaoqing Zheng, Perkin Elmer Centurion Mechanical Seals ...................................................... 165 Non-Contacting Finger Seal Developments and Design Considerations: Thermofluid and Dynamics Characterization, Experimental

M. Jack Braun, Hazel M. Pierson, and Dingeng Deng, University of Akron; and Margaret P. Proctor, NASA Glenn Research Center ................................................................... 181

Role of Distributed Inter-Bristle Friction Force on Brush Seal Hysteresis Helen Zhao and Robert Stango, Marquette University................................................................ 209 DOE/PG&E LNG-Turboexpander Seal and Bearing Retrofit

Donald E. Bently, Dean W. Mathis, and G. Richard Thomas, Bently Pressurized Bearing Company ........................................................................................................................ 223

Advanced Docking/Berthing System Brandan Robertson, NASA Johnson Space Center ..................................................................... 247

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Evaluation of Ceramic Wafer Seals for Future Space Vehicle Applications

Patrick H. Dunlap, Jr., and Bruce M. Steinetz, NASA Glenn Research Center; and Jeffrey J. DeMange, University of Toledo................................................................................... 263

On the Development of a Unique Arc Jet Test Apparatus for Control Surface Seal Evaluations

Joshua R. Finkbeiner, Patrick H. Dunlap, and Bruce M. Steinetz, NASA Glenn Research Center; and Malcolm Robbie, Gus Baker, Arthur Erker, and Joe Assion, Analex Corporation...................................................................................................................... 279

Investigations of High-Temperature Knitted Spring Tubes for Structural Seal Applications

Shawn C. Taylor, Case Western Reserve University; Jeffrey J. DeMange, University of Toledo; and Patrick H. Dunlap, Jr., and Bruce M. Steinetz, NASA Glenn Research Center ...... 297

Modeling of Canted Coil Springs and Knitted Spring Tubes as High-Temperature Seal Preload Devices

Jay J. Oswald and Robert L. Mullen, Case Western Reserve University; and Patrick H. Dunlap, Jr., and Bruce M. Steinetz, NASA Glenn Research Center ........................................... 323

High-Temperature Metallic Seal Development for Aero Propulsion and Gas Turbine Applications

Greg More, Advanced Products; and Amit Datta, Advanced Components and Materials .......... 341 Oxidation of High-Temperature Alloy Wires for Hybrid Seal Applications

Elizabeth J. Opila, NASA Glenn Research Center; Jonathan A. Lorincz, CON/SPAN; Marissa M. Reigel, Colorado School of Mines; and Jeffrey J. Demange, University of Toledo...................................................................................................................................... 359

Attendees List ........................................................................................................................................... 393

OVERVIEW OF NASA GLENN SEAL DEVELOPMENTS

Bruce M. Steinetz, Margaret P. Proctor, and Patrick H. Dunlap, Jr. National Aeronautics and Space Administration

Glenn Research Center Cleveland, Ohio

Irebert Delgado

U.S. Army Research Laboratory Glenn Research Center

Cleveland, Ohio

Joshua Finkbeiner National Aeronautics and Space Administration


Jeffrey J. DeMange and Christopher C. Daniels

University of Toledo Toledo, Ohio

Scott B. Lattime

Ohio Aerospace Institute Brook Park, Ohio

NASA Glenn Research CenterSeal Team

Overview ofNASA Glenn Seal Developments

Dr. Bruce M. SteinetzNASA Glenn Research Center

Cleveland, Ohio

ContributorsMargaret Proctor, Patrick Dunlap, Irebert Delgado

Josh Finkbeiner, Jeff DeMange, Chris Daniels, Scott Lattime

2004 NASA Seal/Secondary Air System Workshop November 9-10, 2004

NASA Glenn Research CenterOhio Aerospace Institute Auditorium

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Overview ofNASA Glenn Seal Developments

Dr. Bruce M. SteinetzNASA Glenn Research Center

Cleveland, Ohio

ContributorsMargaret Proctor, Patrick Dunlap, Irebert Delgado

Josh Finkbeiner, Jeff DeMange, Chris Daniels, Scott Lattime

2004 NASA Seal/Secondary Air System Workshop November 9-10, 2004

NASA Glenn Research CenterOhio Aerospace Institute Auditorium

NASA Glenn hosted the Seals/Secondary Air System Workshop on November 9-10, 2004. At this workshop NASA and our industry and university partners shared their respective seal technology developments. We use these workshops as a technical forum to exchange recent advancements and “lessons-learned” in advancing seal technology and solving problems of common interest. As in the past we are publishing the presentations from this workshop in two volumes. Volume I will be publicly available and individual papers will be made available on-line through the web page address listed at the end of this chapter. Volume II will be restricted under International Traffic and Arms Regulations (I.T.A.R.).

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Workshop AgendaTuesday, Nov. 9, Morning

Registration 8:00 a.m.–8:30 a.m.

Introductions 8:30-9:30 Introduction Dr. Bruce Steinetz, R. Hendricks/NASA GRCWelcome Dr. Rich Christiansen, Deputy Director/NASA GRCOverview of NASA Glenn Seal Program Dr. Bruce Steinetz/NASA GRC

Program Overviews and Requirements 9:30-10:30Overview of NASA’s Exploration Initiative Mr. Joe Naininger for Harry Cikanek/NASA GRCOverview of NASA’s UEET/QAT Project Ms. Carol Ginty/NASA GRC

Break 10:30 -10:45

Turbine Seal Development Session I 10:45-12:30GE90 Aspirating Seal Engine Demonstration Test Ms. Marcia Boyle, B. Albers/GE Aircraft EnginesGeared Fan High Misalignment Seal Development Mr. Dennis Shaughnessy, L. Dobek/Pratt & WhitneyFace Seal Development Mr. Bud Watts for John Munson/Rolls-Royce-AllisonNASA GRC’s Turbine Seal Test Rig: Mr. Irebert Delgado/U.S. Army Res. Lab,

Unique Features M. Proctor/NASA GRCLeakage and Power Loss Test Results for Ms. Margaret Proctor/NASA GRC, I. Delgado/ARL-VTC

Competing Turbine Engine Seals

Lunch: OAI Sun Room 12:30-1:30

The first day of presentations included overviews of NASA programs devoted to the President’s new Space Exploration Initiative and advancing the state-of-the-art in turbine engine technology. Ms. Ginty presented an overview of the Ultra-Efficient-Engine Technology (UEET) and Quiet Aircraft Technology (QAT) programs. The UEET program is aimed at developing highly-loaded, ultra-efficient engines that also have low emissions (NOx, unburned hydrocarbons, etc.). Mr. Naininger of NASA’s Exploration Project office summarized key elements and long range plans for exploring the Moon and Mars through both robotic and manned missions.

Dr. Steinetz presented an overview of NASA seal developments. Representatives from GE provided insight into their advanced seal developments for both aircraft engines and ground power. Mr. Shaughnessy presented an overview of the work P&W and Stein Seal are doing on the development of high misalignment carbon seals for a geared fan application. Mr. Delgado of NASA Glenn presented an overview of turbine testing at NASA GRC. Ms. Proctor presented a comparison of leakage and power loss results for brush and finger seals NASA GRC obtained using the turbine seal rig.

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Workshop AgendaTuesday, Nov. 9, Afternoon

Turbine Seal Development Session II: 1:30-3:30 Tip Clearance

Benefits of High Pressure Turbine Active Clearance Control Mr. Rafael Ruiz, B. Albers/General Electric Aircraft EnginesAADC/Rolls-Royce Active Tip Clearance Control Dev. Mr. Bud Watts, D. Dierksmeier/Allison Advanced Development Co.

-Rolls RoyceMicrowave Blade Tip Sensor Development: An Update Mr. Jon Geisheimer/Radatech Inc.Test Rig for Active Turbine Blade Tip Clearance Dr. Scott Lattime/OAI, B. Steinetz, K. Melcher/NASA GRC,

Control Concepts: An Update J. Decastro/QSSLatest Developments in Wear Prediction of Strip Seals Mr. Norm Turnquist, F. Ghasripoor/GE Global Research Center,

through Conductance M. Kowalczyk, B. Couture/GE Energy

Break 3:30-3:45

Turbine Seal Development Session III 3:45-5:00Parametrical Study of Hydrodynamic Seal Using a 2D Dr. Xiaoqing Zheng/Perkin Elmer Centurion Mech. Seals

Design Code and Comparing with a 3D CFD ModelNon-Contacting Finger Seal Investigations Dr. Jack Braun, H. Pierson, D. Deng, F. Choi/University of AkronNon-Contacting Seal Developments Mr. John Justak/Advanced Technologies GroupRole of Distributed Inter-bristle Friction Force Ms. Helen Zhao, R. Stango/Marquette University

On Brush Seal HysteresisDOE/PG&E LNG-Turboexpander Seal and Bearing Retrofit Dr. Donald Bently, D. Mathis, G. Richard Thomas

Bentley Pressurized Bearing Co.

Group Dinner: Viva Barcelona, Westlake 6:15-?

Turbine engine studies have shown that reducing high pressure turbine (HPT) blade tip clearances will reduce fuel burn, lower emissions, retain exhaust gas temperature margin and increase range. Mr. Ruiz, of General Electric Aircraft Engines, presented results of their Propulsion 21 HPT advanced clearance control study contract. Mr. Watts of Allison Advanced Development Co. presented plans to develop an innovative SMART-Track clearance control mechanism to actively control HPT clearances in the AE30XX series of engines. Mr. Geisheimer of Radatech presented an overview of their microwave blade tip sensor technology. Microwave tip sensors show promise of operation in the extreme gas temperatures present in the HPT location. Dr. Lattime presented the design and development status of a new Active Clearance Control Test rig aimed at demonstrating advanced ACC approaches and sensors. Mr. Turnquist of General Electric Global Research Center presented an overview of wear studies of strip seals used extensively in the ground based power industry.

Dr. Zheng, of PerkinElmer Centurion Mechanical Seals, discussed a non-contacting seal for main shaft locations and component attributes of 2-D and 3-D modeling programs. Dr. Braun presented investigations into a non-contacting finger seal under development by NASA GRC and University of Akron. Mr. Justak presented an overview of non-contacting hybrid seal that combines flexible-beam supported seal pads with a brush secondary seal. Dr. Stango presented analytical assessments of the role of inter-bristle friction force on brush seal hysteresis.

Mr. Richard Thomas of the Bentley Pressurized Bearing Co. presented an overview of a successful program to replace the improperly designed magnetic bearings and brush seals with pressurized bearings and labyrinth seals in a liquid natural gas turboexpander allowing the machine to operate at the required 70,000rpm.

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Workshop AgendaWednesday, Nov. 10, Morning

Registration at OAI 8:00-8:30

Space Vehicle Development 8:30-10:45Future Space Vehicle Docking/Berthing Mechanism and Mr. Brandan Robertson, J. Lewis /NASA JSC

Seal NeedsX-37 Project Overview, Status and Seal Needs Dr. Victor Chen/BoeingOverview of Scramjet Engine Demonstrator Program Mr. Ed Pendleton, AFRL/WPAFB

Break 10:45-11:00

Structural Seal Development Session I 11:00-12:30Evaluation of Ceramic Wafer Seals for Future Mr. Patrick Dunlap, B. Steinetz/NASA GRC,

Space Vehicle Applications J. DeMange/University of ToledoEvaluation of X-37 Flaperon Seal Components Mr. Jeff DeMange/U. of Toledo, P. Dunlap/NASA GRCDevelopment of a Unique Arc Jet Test Apparatus for Mr. Josh Finkbeiner, P. Dunlap, B. Steinetz/NASA GRC

Control Surface Seal Evaluations M. Robbie, A. Erker, J. Assion/AnalexInvestigations of High Temperature Knitted Spring Mr. Shawn Taylor/Case Western Reserve University,

Tubes for Structural Seal Applications J. DeMange/U. of Toledo, P. Dunlap, B. Steinetz/NASA

Lunch OAI Sun Room 12:30-1:30

Mr. Robertson of NASA Johnson Space Center presented an overview of a novel docking and berthing mechanism being developed by NASA JSC with support from NASA Glenn Research Center and Marshall Space Flight Center for future space vehicles as part of NASA’s new Exploration Initiative. This androgynous docking/berthing system would enable any vehicle to dock or berth with any other on-orbit vehicle. To meet this requirement, a seal-on-seal interface is required posing several interesting challenges. Dr. Chen of Boeing-Huntington Beach presented an overview of the X-37 vehicle development status and seal needs. Mr. Ed Pendleton of Wright-Patterson Air Force Base presented a summary of the goals and objectives of the Air Force Scramjet Engine Demonstrator (SED) program.

Mr. Dunlap presented promising flow and high temperature durability results for a ceramic wafer seal being considered for a variety of applications including engine ramps of future hypersonic airbreathing engines. Mr. DeMange presented recent flow and high temperature scrub results for several seals and counter-face materials being considered for the X-37’s control surfaces.

Mr. Finkbeiner presented an overview of an unique arc jet test apparatus being developed to evaluate control surface seals for next generation re-entry and hypersonic vehicles. Mr. Taylor presented high temperature resiliency test results for candidate knitted spring tubes being evaluated for future re-entry vehicle seal needs. Mr. Taylor demonstrated the temperature and resiliency benefits of Rene’41 over the baseline Inconel X-750 wire material.

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Workshop AgendaWednesday, Nov. 10, Afternoon

Structural Seal Development Session II 1:30-3:00Modeling of Canted Coil Springs and Knitted Spring Mr. Jay Oswald, R. Mullen/Case Western Reserve Univ.

Tubes as High Temperature Seal Preload Devices P. Dunlap, B. Steinetz/NASA GRCDevelopment of High Temperature Seal Preloaders: An Update Mr. Ted Paquette/Refractory CompositesHigh Temperature Metallic Seal Development Mr. Greg More/Advanced Products,

A. Datta/Advanced Components & MaterialOxidation of High-Temperature Alloy Wires Ms. Beth Opila/NASA GRC, J. Lorincz/Professional Service

for Hybrid Seal Applications Industries, Inc., M. Reigel/Colorado School of Mines J. DeMange/U. of Toledo

Tour of NASA Seal Test Facilities 3:15-4:15

Adjourn

In the afternoon session, Mr. Oswald presented finite element analysis results for two candidate seal preloaders: the canted coil spring and the knitted spring tube. Mr. Oswald’s results are providing useful insight into the stress states that exist under load helping guide seal preloader design and selection.

Advanced structural seals and preloading elements require application of advanced high temperature materials. The closing session of the workshop presented seal concepts and materials being developed at several locations. Mr. Paquette presented Refractory Composites’ efforts to develop a refractory metal canted coil spring seal preloader, under contract to NASA GRC. Mr. More (Advanced Products) and Dr. Datta (Advanced Components and Materials) presented recent progress in their high temperature (1600-1800°F) metallic seal development.

Another Seal Team goal is to increase the temperature capability of our braided hybrid seal. The hybrid seal combines the features of a lightweight braided ceramic core with an abrasion resistant metallic core outer sheath. Formerly the outer sheath was made of Haynes 188 wires limiting use to <1800°F. Ms. Opila investigated the oxidation behavior of several promising wire materials (e.g. Kanthal and PM2000) for service potentially to 2200°F.

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NASA Glenn Seal Team

Seal Team Leader: Bruce SteinetzMechanical Components Branch/RSM 5950

Turbine Seal DevelopmentDevelop non-contacting, low-leakage turbine sealsMargaret Proctor: Principal Investigator/POCIrebert Delgado, Dave Fleming, Joe Flowers Dan Breen

Structural Seal DevelopmentDevelop resilient, long-life, structural seals for extreme environmentsPat Dunlap: Principal Investigator/POCJeff DeMange, Josh Finkbeiner

Jay Oswald, Shawn TaylorMalcolm Robbie, Art Erker, Joe Assion

Emerging AreasFuel Cell Seals, Acoustic SealsPulse Detonation/Constant Vol. Combustion Engine SealsChris Daniels: Principal Investigator/POCJosh Finkbeiner

Turbine Clearance ManagementDevelop novel approaches for blade-tip clearance control.Scott Lattime: Principal Investigator/POCJim Smialek (5160), Kevin Melcher (5530), Malcolm Robbie

The Seal Team is divided into four primary areas. The principal investigators and supporting researchers for each of the areas are shown in the slide. These areas include turbine seal development, structural seal development, turbine clearance management, and seals for emerging areas. The first area focuses on high temperature, high speed shaft seals for turbine engine secondary air system flow management. The structural seal area focuses on developing resilient structural seals required to accommodate and seal structural distortions in extreme space-and aero-applications. Our goal in the turbine clearance management project is to develop advanced sealing approaches for minimizing blade-tip clearances and leakage. We are planning on applying either rub-avoidance or regeneration clearance control concepts (including smart structures and materials) to promote higher turbine engine efficiency and longer service lives.

We are also contributing seal expertise in a range of emerging areas. These include acoustic seals (a GRC innovation, see Daniels et al, 2004), fuel cell seals, and seals for pulse detonation/constant volume combustion engines. The fuel cell power and pulse detonation engine applications would see significant efficiency gains through the improvement of their sealing systems.

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Attributes of Future Flight Vehicles

Vehicle 50% lighter

25% to 50% more Efficient

PropulsionIntegrated Wing-Body Structure:

25% less drag40% greater range

Extreme Maneuverability and

Control

Highly Intelligent Systems

“Zero” Emissions

“Whisper” Quiet

Attributes of future aircraft are illustrated here. Future vehicles will incorporate advanced materials to reduce weight and drag. Future aircraft will also use highly efficient quiet propulsion systems to reduce fuel burn, reduce emissions and reduce noise in and around airports.

One might ask: What role would advanced seals play in these future vehicles? Lower leakage engine seals reduce engine fuel burn and as a result reduce aircraft emissions. Cycle studies have shown the benefits of increasing engine pressure ratios and cycle temperatures to decrease engine weight and improve performance in next generation turbine engines (Steinetz and Hendricks, 1998). Advanced seals have been identified as critical in meeting engine goals for specific fuel consumption, thrust-to-weight, emissions, durability and operating costs. NASA and the industry are identifying and developing engine and sealing technologies that will result in dramatic improvements and address each of these goals.

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Aspirating Seal Development: GE90 Demo ProgramFunded UEET Seal Development Program

Goal:Complete aspirating seal developmentby conducting full scale (36 in. diameter)aspirating seal demonstration tests inGE90 engine.

Payoffs:- Leakage ~1/4th labyrinth seal - Decrease SFC by 1.86% for three locations- Operates without contact under severe conditions:

+ 10 mil TIR+ 0.25°/0.8 sec tilt maneuver loads

(0.08” deflection!)-

Schedule:– Complete engine assembly: 4Q CY03– GE90 engine test (completed): 1Q CY04

Partners:GE/Stein Seal/CFDRC/NASA GRC

General Electric GE90

Sealing dam

Retraction spring

Hydrostatic bearing

Rotor

Phi

Labyrinth seal

General Electric is developing a low leakage aspirating face seal for a number of locations within modern turbine applications. This seal shows promise both for compressor discharge and low-pressure turbine balance piston locations.

The seal consists of an axially translating mechanical face that seals the face of a high speed rotor (Turnquist et al, 1999). The face rides on a hydrostatic cushion of air supplied through ports on the seal face connected to the high pressure side of the seal. The small clearance (0.001-0.002 in.) between the seal and rotor results in low leakage (1/4th that of new labyrinth seals). Applying the seal to 3 balance piston locations in a GE90 engine can lead to >1.8% SFC reduction. GE Corporate Research and Development tested the seal under a number of conditions to demonstrate the seal’s rotor tracking ability. The seal was able to follow a 0.010 in. rotor face total indicator run-out (TIR) and could dynamically follow a 0.25˚ tilt maneuver (simulating a hard maneuver load) all without face seal contact. The NASA GRC Ultra Efficient Engine Technology (UEET) Program funded GE to demonstrate this seal in a ground-based GE-90 demonstrator engine in 2003-2004. More details can be found in Boyle and Albers, 2005 in this Seal Workshop Proceedings and Turnquist, et al 1999.

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Non-Contacting Finger Seal DevelopmentNASA GRC/University of Akron

Objective:Develop non-contacting finger seal to overcome

finger element wear and heat generation for future turbine engine systems

Approach:• Solid modeling for finger and pad

motion/stresses • Fluid/solid interaction for leakage evaluation• Experimental verification

Status:• Developed a simplified spring-mass-damper

model to assess seal’s dynamic response.• CFD-ACE+ (3-D Navier-Stokes code) utilized

to analyze the thermofluid behavior and to obtain stiffness and damping parameters.

• First prototype built: Testing underway

Program:NASA/Univ. of Akron Coop. Agreement: Dr. Braun (U. of Akron) M. Proctor, Monitor

Axial View of Staggered High/Low Pressure Fingers Assembly (View A-A)

A

A

Downstream Fingers With Pads (Low P)

Upstream Fingers(High P)

Seal Prototype

1 2

Conventional finger seals like brush seals attain low leakage by operating in running contact with the rotor (Proctor, et al, 2002). The drawbacks of contacting seals include wear over time, heat generation, and power loss (Proctor and Delgado, 2004).

NASA Glenn has developed several concepts for a non-contacting finger seal. In one of these concepts the rear (low-pressure, downstream) fingers have lift pads (see pads 1 & 2 in inset figure) and the upstream (high pressure side) fingers are pad-less, and are designed to block the flow through the slots of the downstream fingers. The pressure-balance on the downstream-finger lift-pads cause them to lift. The front fingers are designed to ride slightly above the rotor preventing wear. Pressure acts to hold the upstream fingers against the downstream fingers. It is anticipated that the upstream/downstream fingers will move radially as a system in response to shaft transients. The NASA Glenn non-contacting finger seal was recently awarded a U.S. patent No. 6,811,154. (Proctor and Steinetz, 2004)

Dr. J. Braun of University of Akron is performing analyses and tests of this GRC concept through a cooperative agreement (Braun et al, 2003). University researchers developed an equivalent spring-mass-damper system to assess lift characteristics under dynamic excitation. Fluid stiffness and damping properties were obtained utilizing CFD-ACE+ (3-D Navier-Stokes code) and a perturbation approach. These stiffness and damping properties were input into the dynamic model expediting the solution to aid in the design of the finger and pod configurations. Dr. Braun and his team subsequently fabricated a first generation non-contacting finger seal based on this design. Early test results showed that the finger seal operates without contact with the shaft at pressures up to 15 psid. Non-contact operation was proven via both electric-circuit continuity and high magnification photo-imagery. More details can be found in Braun et al, 2005 in this Seal Workshop Proceedings. After feasibility tests are complete at the University, seals will be tested under high speed and high temperature conditions at NASA GRC.

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Turbine Clearance Management Goal

Develop and demonstrate clearance management technologies to improve turbine engine performance, reduce emissions, and increase service life

HPT blade

HPT disk

CDP air

HPT blade tip seal

Combustor

System studies have shown the benefits of reducing blade tip clearances in modern turbine engines, especially the high pressure turbine. Minimizing blade tip clearances throughout the engine will contribute materially to meeting NASA’s Ultra-Efficient Engine Technology (UEET) turbine engine project goals of reducing fuel burn and emissions. NASA GRC is examining two candidate approaches including rub-avoidance and regeneration.

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Tip Clearance Variation: Motivation for Clearance Management

SOA Thermal Clearance Control

Active Clearance Control

Cruise(new engine)

Cruise (worn engine)

15-20 mil

30-50 mil ~5 mil

~5 mil

0-10 mil ~5 milblade

shroudTakeoff

(new engine)blade

shroud

The Problem:Clearances between the shroud and blade tips vary over the operation and life of an engine. Wear and thermal erosion increases blade tip clearance.

Impacts:• Loss of engine efficiency & increased SFC• Increase in NOx & CO emissions• Rise in exhaust gas temperature (EGT)

ACC System Challenges:

Temperature: Gas path - >2500°FCooling air - >1200°FCase - 600°F (w/ soak back)

Load/Response: Actuators must react ~2000 lbf move ~0.05” in 10 sec

Accuracy: Current Systems - 0.015-0.020-inGoal – <0.005-in

Size/Weight: Small, lightweight ACC systems requiredGoal current thermal systems (<100 lbs).

Clearances between the shroud and blade tips vary over the operation and life of an engine. During operation, variations in tip clearance occur primarily due to differences in thermal growth of the case and thermal-mechanical growth of the rotor. Wear of the shroud and blade tips due to rubs and thermal erosion increases over the engine's life and contributes to permanent increases in blade tip clearance. As clearances increase, the engine runs hotter (less efficient) to achieve the same thrust and speed.

Current ACC systems use fan and compressor air to contract the HPT case flanges, varying the shroud diameter, and hence blade-tip clearance during cruise (see thermal clearance control illustration above). These systems cannot respond to fast transient events such as takeoff, re-accel, and sudden step altitude changes. Adequate cold-build clearances are chosen to prevent rubbing between the blade tips and shroud seals during minimum clearance events. As such, tip clearances are larger than desired throughout the flight profile causing greater fuel burn and emissions and shorter range. Utilizing our proposed fast-response, mechanical systems will minimize clearances throughout flight operation, including fast transient events. (See active clearance control illustration above)

There are a number of challenges that must be overcome in developing a successful active clearance control (ACC) system. One of the largest challenges of those listed above is the extreme thermal environment the ACC system must operate in. Gas path temperatures exceed 2500°F and shroud cooling temperatures exceed 1200°F.

NASA/CP—2005-213655/VOL1 12


• Fuel Savings/ Reduced Emissions (HPT)– 0.010-in tip clearance is worth ~0.8-1% SFC– Emissions Reduction (Landing/Takeoff – Ref. GE

Propulsion 21 Study)»NOx»CO

• Increased Cycle Life (Reduced Maintenance Costs)

– Deterioration of exhaust gas temperature (EGT) margin is the primary reason for aircraft engine removal from service

– 0.010-in tip clearance is worth ~10 ºC EGT– Allows turbine to run at lower temperatures,

increasing cycle life of hot section components and engine time-on-wing (~1000 cycles)

• Increased Efficiency/Operability– Increased payload and mission range capabilities– Increased high pressure compressor (HPC) stall

margin

Benefits of Blade Tip Clearance Control

Clearance Control Technology Promotes High Efficiency and Long Life

Exh

aust

Gas

Tem

per

atu

re(E

GT

) o

r 1/

Eff

icie

ncy

Increased Time-on-Wing

Time

Engine withoutclearance control technology:Shorter service life

FAA EGT Limit

Engine withclearance control technology:Longer service life

Blade tip clearance opening is a primary reason for turbine engines reaching their FAA certified exhaust gas temperature (EGT) limit and subsequent required refurbishment. As depicted in the chart on the right, when the EGT reaches the FAA certified limit, the engine must be removed and refurbished. By implementing advanced clearance control measures, the EGT rises slower (due to smaller clearances) increasing the time-on-wing.

In summary, benefits of clearance control in the HPT turbine section include lower specific fuel consumption (SFC), lower emissions (NOx, CO), retained exhaust gas temperature (EGT) margins, higher efficiencies, longer range (because of lower fuel-burn) (see General Electric Report, 2004 and Lattime et al, 2002). Benefits of clearance control in the compressor include better compressor stability (e.g. resisting stall/surge), higher stage efficiency, and higher stage loading. All of these features are key for future NASA and military engine programs.

NASA/CP—2005-213655/VOL1 13


Active Clearance Control Concept & Evaluation Test Rig

Heat Inputs:+ Radiant+ Air Supply

Chamber

Seal carrier assembly

ActuatorsGen 1: servo-hydraulic

Purpose:• Evaluate actuator concept

response and accuracy under appropriate thermal (to 1500°F) and pressure (to 120 psi) conditions.

• Evaluate clearance sensor response and accuracy

– Capacitance– Microwave

• Measure secondary seal leakage due to segmented design and case penetration.

AdvancedClearance Sensors

NASA GRC is developing a unique Active Clearance Control (ACC) concept and evaluation test rig. The primary purpose of the test rig is to evaluate actuator concept response and accuracy under appropriate thermal (to 1300+F) and pressure (up to 120 psig) conditions. Other factors that will be investigated include:

•Actuator stroke, rate, accuracy, and repeatability

•System concentricity and synchronicity

•Component wear

•Secondary seal leakage

•Clearance sensor response and accuracy

The results of this testing will be used to further develop/refine the current actuator design as well as other advanced actuator concepts. More details regarding this test rig can be found in Lattime and Steinetz 2005 in this Seal Workshop Proceedings, and Lattime et al, 2003.

NASA/CP—2005-213655/VOL1 14


ACC Test Rig Details

Multiple independently controlled actuators permit either axisymmetric or asymmetric control

radiant heater

inlet air (Phigh)

exhaust air (Plow)

chamber

seal carrier

proximity probe

footactuator rod

main housing

chamber support tubeactuator

movement

chamber metal TC’s

chamber air TC

flow deflector

actuator mount

radiant heater

inlet air (Phigh)

exhaust air (Plow)

chamber

seal carrier

proximity probe

footactuator rod

main housing


movement


chamber air TC

flow deflector

actuator mount

Axisymmetric Asymmetric

turbine wheelturbine shroud

turbine wheelturbine shroud

ACC Test Rig fabrication complete(rig assembly at vendor)

The ACC test rig under development utilizes nine independently controlled actuators. This will permit assessment of both axisymmetric control and asymmetric control. Because of engine thermal and structural non-uniformities, engine case structures can become egg-shaped. Under these circumstances, asymmetric control strategies would permit more uniform clearances around the circumference.

The ACC test rig fabrication has been completed, as shown in the lower right figure. However, the hydrotest performed revealed some weld cracks that are currently being investigated.

NASA/CP—2005-213655/VOL1 15


NASA GRC Structural Seal Development Goals

Develop hot (2000-2500+°F), flexible, dynamic structural seals for ram/scramjet propulsion systems (TBCC, RBCC)

Develop reusable re-entry vehicle control surface seals to prevent ingestion of hot (6000 °F) boundary layer flow

Example: X-37; X-38 CRV

ControlSurface Seals

ISTAR EngineInlet Ramp Seal

LRC

TBCCRam/Scramjet Engines

RBCC

NASA GRC is developing advanced structural seals for both propulsion and vehicle needs by applying advanced design concepts made from emerging high temperature ceramic materials and testing them in advanced test rigs. See Dunlap 2005, et al, and Finkbeiner 2005, et al in this Seal Workshop Proceedings and Dunlap 2004a, b, et al and DeMange 2004, et al for further details.

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FALCON Hypersonic Vehicle Seal Development

• Objective: Develop high temperature seals for control surfaces and access doors on future hypersonic vehicles

• Requirements– Temperature: Extreme– Life: Reusable– Mission duration: Less than 2 hrs

• Approach– Identify and develop high

temperature seals and preload devices

– Perform critical function performance tests at GRC

– Perform arc jet tests on leading concepts at JSC

• Partner organizations: DARPA, Lockheed Martin

Model of GRC seal arc jet test fixture

NASA GRC is working under DARPA sponsorship to develop control surface seals for future hypersonic weapon systems under a program known as FALCON.

In this program we plan on identifying and developing seals and preloader system that can meet the expected temperatures and pressures. We will perform critical function performance tests utilizing our new test capabilities described on the next chart. Those concepts that meet functional requirements will then be tested in an arc jet environment at NASA JSC. We will use an arc jet test fixture that can subject candidate seals and ceramic matrix composites to simulated hypersonic flight conditions.

A unique feature of this new test fixture is the ability to asses the effects of flap motion on seal performance during arc jet testing. More detail regarding this fixture can be found in Finkbeiner, et al, 2004 and Finkbeiner, et al 2005 in this Seal Workshop Proceedings.

NASA/CP—2005-213655/VOL1 17


Example Structural Seals Being Investigated

Ceramic Wafer Seal• High temperature operation: 2500+˚F• Low Leakage• Flexibility: Relative sliding of adjacent wafers

conforms to wall distortions• Ceramic material lighter weight than metal

system• Tandem seals permit central cavity purge

(cooling)

Braided Rope Seal• High temperature operation: 2400+°F• Flexible: seals & conforms to complex

geometries• Hybrid design (ceramic core/superalloy wire

sheath) resists abrasion• Tandem seals permit central cavity purge

(cooling)

NASA GRC’s work on high temperature structural seal development began in the late 1980’s during the National Aero-Space Plane (NASP) project. GRC led the in-house propulsion system seal development program and oversaw industry efforts for propulsion system and airframe seal development for this vehicle.

Two promising concepts identified during that program included the ceramic wafer seal (Dunlap, 2004a et al and Steinetz, 1991) and the braided rope seal (DeMange, 2004 et al, Steinetz and Adams, 1998) shown here. By design, both of these seals are flexible, lightweight, and can operate to very high temperatures (2400+˚F). Both types of seals require some form of high temperature preload system. A high temperature canted coil spring is shown behind the seals shown. More information on the features and benefits of the canted coil spring can be found in Dunlap et al, 2004b and Oswald et al 2004 and Oswald 2005, et al in this Seal Workshop Proceedings. Refractory metals and oxygen resistant coatings being considered for the spring can be found in Paquette, 2005 in this Seal Workshop Proceedings.

A second type of preload system is also under development at GRC for textile based thermal barriers. A spring tube knitted out of Rene ’41 shows greater resiliency at both 1500° and 1750°F than if made of the conventional Inconel X-750 material. (Taylor et al, 2004 and in Taylor et al, 2005 in this Seal Workshop Proceedings.)

NASA/CP—2005-213655/VOL1 18


Hot Compression/Scrub Seal Test Rig: Overview

Load frame

Laser extensometer

3000 °F furnace

Seal

Seal holder

Wafer seals

Seal holder

Inconel or Shuttle tile rub surfaces

NASA GRC has installed state-of-the-art test capabilities for evaluating seal performance at temperatures up to 3000 °F (1650 °C). This one-of-a-kind equipment is being used to evaluate existing and new seal designs by simulating the temperatures, loads, and scrubbing conditions that the seals have to endure during service. The compression test rig (upper left photo) is being used to assess seal load vs. linear compression, preload, & stiffness at temperature. The scrub test rig (middle photo) is being used to assess seal wear rates and frictional loads for various test conditions at temperature. Both sets of fixtures are made of silicon carbide permitting high temperature operation in air.

The test rig includes: an MTS servo-hydraulic load frame, an ATS high temperature air furnace, and a Beta LaserMike non-contact laser extensometer, and the special purpose seal holder hardware. Unique features of the load frame include dual load cells (with multi-ranging capabilities) for accurate measurement of load application, dual servo-valves to permit precise testing at multiple stroke rates (up to 8 in./s.), and a non-contact laser extensometer system to accurately measure displacements.

NASA/CP—2005-213655/VOL1 19


Shuttle Main Landing Gear Door Seal Tests

• Objective: Perform flow and compression tests on Shuttle main landing gear (MLG) door environmental seal at different compression levels

• Why important?– JSC using data to determine acceptable seal

gap range that can be verified each time MLG doors are closed for flight

– Concerned about long term seal creep; need way to predict when to change out seals

• Two phases of testing– Phase I (Summer 2004): Flow tests on “as-

received” seals– Phase II (1Q FY05):

» Flow tests on “flown” seals» Compression tests on “flown” and “as-

received” seals including 30-day compression test

At the request of NASA Johnson Space Center (JSC), a series of tests are being conducted by the Seal Team at GRC on the Shuttle main landing gear (MLG) door environmental seal in support of the Shuttle Return-to-Flight Program. This includes both flow tests and compression tests on the seals at different compression levels. JSC is using the data to determine an acceptable seal gap range that can be verified each time the MLG doors are closed for flight. They are also concerned about the effects of long-term creep when the seal is compressed for extended periods of time. The seal takes on a permanent set under these conditions and may not stay in contact with the opposing sealing surface at all times. JSC would like to determine a way to predict when the seals need to be replaced to avoid possibly dangerous situations that could occur if they take on too much permanent set.

This work has been divided into two phases of testing. This past summer the Seal Team completed a series of flow tests for JSC on samples of the “as-received” seal that had not yet flown on the Shuttle. Upon recently inspecting the MLG door environmental seals installed on several of the Shuttle orbiters, JSC discovered that they had taken on a set and were permanently compressed. JSC asked the Seal Team to repeat the sequence of leakage tests on these “flown” seals to determine how much flow gets past them if the seals do not stay in contact with the sealing surfaces in their permanently-compressed state. JSC has also requested that a series of compression tests be performed on as-received and flown seals to determine how much resiliency they have in each condition. A 30-day compression test will also be performed to evaluate the effects of long-term seal creep on seal resiliency.

NASA/CP—2005-213655/VOL1 20


Shuttle MLG Door Seal Tests – Phase I Results

• Performed series of flow tests on “as-received” seals at different compression levels and gap sizes

• Seal: Molded silicone rubber tadpole wrapped with Nomex fabric

• Amount of flow past seals decreased as:

– Amount of compression increased

– Gap size decreased

• Shared results with JSC to input into their models

• Tests will be repeated using “flown” seals in Phase II tests

Main landing gear door environmental seal

Flow Results

0.0

0.4

0.8

1.2

1.6

2.0

0 3 6 9 12 15

Delta P across seal (psid)

Flo

w p

er in

ch o

f se

al

(SL

PM

/in)

0.410-in. gap0.345-in. gap0.270-in. gap0.195-in. gap0.120-in. gap

Vent hole

Hollow bulb

Attachment tail

Sealing surface

In Phase I, a series of successful flow tests were conducted on the “as-received”seals at different compression levels and gap sizes. Samples were leak tested at pressures up to 14.7 psid. This seal is currently produced by Northrop Grumman and consists of a silicone (ZZ-R-765, Class IIIa, Grade 50) core and tail which are wrapped with Nomex fabric. As expected, the amount of flow past the seals decreased as the amount of compression on the seals was increased and the gap size decreased. A final report was completed and forwarded to JSC for review and further analysis. These flow tests will be repeated in Phase II using “flown” seals that were removed from the Shuttle.

NASA/CP—2005-213655/VOL1 21


Shuttle RCC Leading Edge Permeability Tests

• Objective: Measure permeability of Shuttle leading edge reinforced carbon/carbon (RCC) material (coated & uncoated) in support of Return-to-Flight program

• Why important?– Material permeability determines amount of hot gas that can pass through leading edge

and into cavity behind it

– JSC using permeability values for carbon/carbon oxidation calculations under reentry conditions

Individual leading edge panel

The GRC Seal Team is also supporting the Return-to-Flight program by measuring the permeability of the reinforced carbon/carbon (RCC) material that makes up the Shuttle leading edges. This material is being evaluated in both coated and uncoated states to determine how the coating affects RCC permeability. These tests are important because the permeability of this material determines how much hot gas is able to pass through the leading edge and into the cavity behind it where lower-temperatures structures are located. JSC is also planning to use the permeability values that GRC records to calculate oxidation of the RCC material under reentry conditions.

NASA/CP—2005-213655/VOL1 22


Shuttle RCC Leading Edge Permeability Tests

• New test fixture used to measure permeability of porous and semi-porous materials based on data recorded during leak decay testing

• Automated data system tracks pressure decay vs. time

• Data used to calculate mass flow and permeability

• Initial results for coated RCC specimen show good correlation in mass flow vs. ΔP between GRC data and data recorded by Vought in 1980

Pressure transducersPressurized cylinder in water

bath (for temp. control)

Mass Flow Rate vs Pressure Drop

0.0E+00

3.0E-06

6.0E-06

9.0E-06

1.2E-05

1.5E-05

1.8E-05

0 1 2 3 4 5 6 7

Specimen Pressure Drop [psi]

Mas

s F

low

Rat

e [l

bm

/s] GRC Data

GRC RepeatGRC Reassemble, RepeatVought DataVought Data, No Glaze

Specimen (2”x2”) in

holder

To perform these tests, GRC has installed a new test fixture that is being used to measure material permeability based on data recorded during leak decay tests. During these tests, a cylinder upstream of the test specimen is pressurized to a desired pressure. This pressure is then allowed to decay as air flows through the test specimen. An automated data acquisition system monitors the pressure decay versus time, and this data is then used to calculate the amount of mass flow through the specimen and the permeability of the material. Initial results for a coated RCC specimen have shown good correlation in the amount of mass flow through the specimen versus the pressure drop across the specimen when comparing GRC data to data presented in a report by Vought in 1980. The next step is to test an uncoated specimen and compare its permeability to that of the coated specimen.

NASA/CP—2005-213655/VOL1 23


Seal Development for Advanced Docking/Berthing SystemSpace Exploration Initiative

• Objective: Support NASA JSC by developing seals for advanced docking/berthing system

• Requirements– Seal diameter: 54 in.– Near hermetic seal– Seal-on-seal interface for androgynous system– Survive space environment (atomic oxygen, UV,

thermal cycling) for long duration (est. 3-5+ yrs)

• Approach– Identify candidate elastomeric and metallic seals– Perform coupon-level and small-scale

environmental exposure and flow tests– Perform mid-scale flow tests before and after

environmental exposure (incl. variable gap, offset, hot and cold, seal-on-seal)

– Down-select between competing concepts– Support scale-up and thermal-vacuum system-

level tests at MSFCSeal-on-seal interface

Docking/BerthingSystem

NASA plans on developing an advanced docking and berthing system that would permit any vehicle to dock to any on-orbit station or vehicle, as part of NASA’s new Exploration Initiative. (More detail on this new docking and berthing system can be found in Robertson 2005, in this Seal Workshop Proceedings.) To meet this “androgynous” operational goal, a seal-on-seal interface is required, as depicted in the lower graphic.

GRC will be supporting JSC in developing seal technology for this seal interface. Any seal developed must meet the stringent requirements identified including near-hermetic operation, prevent seal pull-out and resist space environments (atomic oxygen, UV, radiation and thermal cycling) for five plus years, amongst others. An evolutionary development approach has been identified as outlined above.

NASA/CP—2005-213655/VOL1 24


Solid Oxide Fuel Cells and NASA Applications

Auxiliary Power Units (APU)• Greater efficiency than traditional

turbine APUs• Aviation fuel capable and facilitates

transition to H2 based systems• Up to 20% Ground / Landing-Take-

Off aircraft NOx reduction• About 20 dB noise reduction at the

gateNet reaction: 2H2 + O2 → 2H2O

Electrolyzer• Night time power-production• “Reversed” cell operation can create

– Breathable Oxygen– Fuel for return-to-Earth

Unmanned Aerial Vehicles• Increased fuel economy

– More time aloft for mission• Emissions reduction

How Does it Work?

Fuel Cell APU

Turbine APU

A fuel cell is an energy conversion device that generates electricity and heat by electrochemically combining fuel and oxidizer via an ion-conducting electrolyte. The electrochemical reaction is more attractive than combustion since the process is more efficient and less polluting. Solid Oxide Fuel Cells (SOFCs) are unique in that they are fuel flexible, using hydrogen, carbon monoxide, natural or coal gas, or low-sulfur jet fuel as the reactant.

NASA has three potential applications for SOFCs... (1) Commercial aircraft would benefit greatly by the substitution of SOFC-powered auxiliary power units (APUs) with (a) increased efficiency at both ground idle and in cruise, (b) reduced emissions of NOx, SOx, and particulates, (c) potentially greater power generating capability at altitude, and (d) gate noise reduction. (2) A hydrogen powered unmanned aerial vehicle (UAV) capable of flying for up to 10 days and solar-electric aircraft capable of multiple-week duration flights are both candidates for SOFC propulsion power. (3) A reversible SOFC could provide Space Exploration Initiative missions with power generation for base operations, fuel generation for the return trip, and breathable oxygen for astronaut explorers.

NASA/CP—2005-213655/VOL1 25


Fuel Cell Seal Development

What are the problems with SOFC seals?

• Coefficient of Thermal Expansion (CTE) mismatch between the adjacent components causes relative displacements and seal damage

• Loss of seal integrity reduces SOFC performance due to

– Fuel/air leakage– Electrode poisoning

Approach: Pursue a multidisciplinary development effort including:

• Thermo-structural analyses• Novel seal design concepts • Advanced materials• Experimentation

Test Apparatus Capabilities:– Temperatures (up to 1100°C)– Temperature transients (approx.

1°C/second)

– Gases (helium, air, and other non-combustibles)

– Gas pressures (up to 2.5 psig)

– Mechanical loads (up to 1000 lbf)

Seal

Gas Supply

Ceramic/ Metal Sealing Surfaces

Planar SOFCs require high temperature hermetic seals to (1) prevent mixing of the fuel and oxidant within the stack, (2) prevent parasitic leakage of the fuel from the stack, (3) prevent contamination of the anode by air leaking into the stack, (4) electrically isolate the individual cells within the stack, and (5) mechanically bond the cell components. The sealing challenges are aggravated by the need to maintain hermetic boundaries between the different flow paths during transient (heating / cooling) operation with vibration loads.

NASA GRC is taking a multidisciplinary approach to developing SOFC seals. Thermo-structural analyses, novel seal concept development, advanced materials development, and experimental leakage determination are being simultaneously pursued to solve the sealing challenges.

A leakage facility has been built at NASA GRC to accurately measure any leakage through SOFC seals while simultaneously exposing the seal to an environment that closely resembles that of an operating fuel cell. Capable of heating to temperatures as high as 1100 Celsius at ramp rates of approximately 1 degree Celsius per second, the system measures the leakage of any non-combustible gas while maintaining compressive mechanical loads on the seal.

NASA/CP—2005-213655/VOL1 26


• Seals technology recognized as critical in meeting next generation aero- and space propulsion, power and space vehicle system goals

• Performance • Efficiency • Life/Reusability• Safety• Cost

• NASA Glenn is developing seal technology and/or providing technical consultation for the Nation’s key aero- and space advanced technology development programs.

Summary

NASA Glenn is currently performing seal research supporting both advanced turbine engine development and advanced space vehicle/propulsion system development. Studies have shown that decreasing parasitic leakage through applying advanced seals will increase turbine engine performance and decrease operating costs.

Studies have also shown that higher temperature, long life seals are critical in meeting next generation space vehicle and propulsion system goals in the areas of performance, reusability, safety, and cost.

NASA Glenn is developing seal technology and providing technical consultation for the Agency’s key aero- and space technology development programs.

NASA/CP—2005-213655/VOL1 27


NASA Seals Web Sites

• Turbine Seal Development– http:/www.grc.nasa.gov/WWW/TurbineSeal/TurbineSeal.html

» NASA Technical Papers

» Workshop Proceedings

• Structural Seal Development– http://www/grc.nasa.gov/WWW/structuralseal/

» NASA Technical Papers

» Discussion» Seal Patents

– http://www/lerc.nasa.gov/WWW/TU/InventYr/1996Inv_Yr.htm

The Seal Team maintains three web pages to disseminate publicly available information in the areas of turbine engine and structural seal development. Please visit these web sites to obtain past workshop proceedings and copies of NASA technical papers and patents.

NASA/CP—2005-213655/VOL1 28


References

• Braun, M.J., Choy, F.K., Pierson, H.M., 2003, “Structural and Dynamic Considerations Towards the Design of Padded Finger Seals”, AIAA-2003-4698 presented at the AIAA/ASME/SAE/ASEE conference, July, Huntsville, AL.

• Daniels, C., Finkbeiner, J, Steinetz, B.M., Li, Xiaofan, Raman, G, 2004, “Non-linear Oscillations and Flow of Gas Within Closed and Open Conical Resonators,” AIAA-2004-0677. To be presented at the 82nd AIAA Aerosciences Meeting, Reno, NV, January.

• DeMange, J.J., Dunlap, P.H., Steinetz, B.M., 2004,“Advanced Control Surface Seal Development for Future Space Vehicles,” Presentation and Paper at 2003 JANNAF Conference, Dec. 1-5, Colorado Springs, CO. (NASA TM-2004-212898).

• Dunlap, P.H., Steinetz, B.M., DeMange, J.J., 2004a, “Further Investigations of Hypersonic Engine Seals,” NASA TM-2004-213188, AIAA-2004-3887, August. Presented at the 2004 AIAA/ASME/SAE/ASEE Joint Propulsion Conference, July 2004, Ft. Lauderdale, FL.

• Dunlap, P.H., Steinetz, B.M., DeMange, J.J., 2004b, “High Temperature Propulsion System Structural Seals for Future Space Launch Vehicles,” Presentation and Paper at 2003 JANNAF Conference, Dec. 1-5, Colorado Springs, CO. (NASA TM-2004-212907).

• Finkbeiner, J.R., Dunlap, P.H., Steinetz, B.M., Robbie, M., Baker, F., and Erker, A., 2004, “On the Development of a Unique Arc Jet Test Apparatus for Control Surface Seal Evaluations,” NASA TM-2004-213204, AIAA-2004-3891, August. Presented at the 2004 AIAA/ASME/SAE/ASEE Joint Propulsion Conference, July 2004, Ft. Lauderdale, FL.

• General Electric Aircraft Engines, 2004, “HPT Clearance Control (Intelligent Engine Systems) – Phase 1 – Final Report,” NASA Contract NAS3-01135, April.

• Lattime, S.B., Steinetz, Bruce M., Robbie, M., 2003, “Test Rig for Evaluating Active Turbine Blade Tip Clearance Control Concepts,” NASA TM-2003-212533, also AIAA-2003-4700, presented at the AIAA/ASME/SAE/ASEE conference, July, Journal of Propulsion and Power, 2004, Huntsville, AL.

• Lattime, S.B., Steinetz, B.M., 2002, “Turbine Engine Clearance Control Systems: Current Practices and Future Directions, “ NASA TM-2002-211794, AIAA 2002-3790.

NASA/CP—2005-213655/VOL1 29


• Oswald, J.J., Mullen, R.H., Dunlap, P.H., and Steinetz, B.M., 2004, “Modeling and Evaluation of High Temperature Canted Coil Springs as Seal Preloading Devices,” NASA TM-2004-213189, AIAA-2004-3889, August. Presented at the 2004 AIAA/ASME/SAE/ASEE Joint Propulsion Conference, July 2004, Ft Lauderdale, FL.

• Proctor, M.P., Delgado, I.R., 2004, “Leakage and Power Loss Test Results for Competing Turbine Engine Seals,” ASME GT2004-53935, Proceedings of ASME Turbo Expo, Vienna Austria, June.

• Proctor, M.P., Steinetz, B.M., 2004, “NonContacting Finger Seal,” U.S. Patent 6,811,154, Issued 11/02/04, (LEW 17,129-1).

• Proctor, M.P; Kumar, A.; Delgado, I.R., 2002, “High-Speed, High Temperature, Finger Seal Test Results,” NASA TM-2002-211589, AIAA-2002-3793.

• Steinetz, B.M., Hendricks, R.C., and Munson, J.H., 1998, “Advanced Seal Technology Role in Meeting Next Generation Turbine Engine Goals,” NASA TM-1998-206961.

• Steinetz, Bruce M.; Adams, Michael L., 1998, “Effects of Compression, Staging and Braid Angle on Braided Rope Seal Performance”, J. of Propulsion and Power, Vol. 14, No. 6, also AIAA-97-2872, 1997 AIAA Joint Propulsion Conference, Seattle, Washington, July 7-9, 1997, NASA TM-107504, July 1997.

• Steinetz, B.M., 1991, “High Temperature Performance Evaluation of a Hypersonic Engine Ceramic Wafer Seal,” NASA TM-103737

• Taylor, S.C., DeMange, J.J., Dunlap, P.H., Steinetz, B.M., 2004, “Evaluation of High Temperature Knitted Spring Tubes for Structural Seal Applications,” NASA TM-2004-213183, AIAA-2004-3890. Presented at the 2004 AIAA/ASME/SAE/ASEE conference, July, Ft Lauderdale, FL.

• Turnquist, N.A.; Bagepalli, B; Lawen, J; Tseng, T., McNickle, A.D., Kirkes; Steinetz, B.M., 1999, “Full Scale Testing of an Aspirating Face Seal”, AIAA-99-2682.

References (cont’d)

NASA/CP—2005-213655/VOL1 30

OVERVIEW OF NASA’S EXPLORATION INITIATIVE

Joseph Nainiger National Aeronautics and Space Administration


Exploration Systems Mission Directorate

Overview of NASA’s

Exploration Initiative

Joseph J. NainigerDeputy Chief for CapabilityExploration Systems DivisionNASA GRC, Cleveland, OhioNovember 9, 2004

NASA/CP—2005-213655/VOL1 31

The Vision for Space Exploration

Implement a sustained and affordable human and robotic program to explore the solar system and beyond

Extend human presence across the solar system, starting with a human return to the Moon by the year 2020, in preparation for human exploration of Mars and other destinations;

Develop the innovative technologies, knowledge, and infrastructures both to explore and to support decisions about the destinations for human exploration; and

Promote international and commercial participation in exploration to further U.S. scientific, security, and economic interests.

THE FUNDAMENTAL GOAL OF THIS VISION IS TO ADVANCE U.S. SCIENTIFIC, SECURITY, AND ECONOMIC INTEREST THROUGH A ROBUST

SPACE EXPLORATION PROGRAM

• On January 14th, the President announced that we were going back to the moon, continuing to Mars and Beyond.

• The plan to accomplish this is laid out in the Vision document

• Read “Fundamental Goal” Statement on chart

• Scientific – From a scientific standpoint, exploration and science leads to discovery, which is what drives us

• Security – Security in this sentence is not homeland security, but rather allowing other gov’t agencies to leverage our technologies to further their objectives.

• Economic Interest - For less than 1% of the budget, we have improved the quality of life and the aerospace industry has a major economic impact on the nation. We know we will have the same type of return through exploration and discovery in the future, we just can’t define what specifically it will be.

• The key word here is sustainability – we are in it for the long haul – we have to get through 30 budget cycles, 8 presidential elections and multiple congresses. We need to be credible. When you have credibility and affordability, you have sustainability.

NASA/CP—2005-213655/VOL1 32

• Objectives

– Implement a sustained and affordable human and robotic program

– Extend human presence across the solar system and beyond

– Develop supporting innovative technologies, knowledge, and infrastructures

– Promote international and commercial participation in exploration

• Major Milestones

– 2008: Initial flight test of CEV

– 2008: Launch first lunar robotic orbiter

– 2009-2010: Robotic mission to lunar surface

– 2011 First Uncrewed CEV flight

– 2014: First crewed CEV flight

– 2015: Prometheus 1 - Jupiter Icy Moon Orbiter (JIMO)

– 2015-2020: First human mission to the Moon

Key Elements of the Vision

➘ We have identified a number of transforming changes discussed in our Strategic Plan that are important ingredients for fulfilling our Vision & Mission

➘ Highlight SUSTAINED being through the future not just for set amount of years.

➘ Establish credibility by achieving these milestones

➘ On time and within budget

➘ Sept 04 a RFI was sent out and over 1000 concepts were received for CEV and 11 selections have been made

➘ Talk to timeline (major milestones)

➘ 2012 -2015 Maintaining the importance of Nuclear propulsion research through JIMO

Other questions if asked:

Difference from OSP to CEV is that OSP was low orbit crew return vehicle to ISS, but CEV is planned to travel to the moon and mars.

NASA/CP—2005-213655/VOL1 33

Realizing the FutureEarth, Moon, Mars, and Beyond

Foster and sustain the exploration culture across generations

• Open new frontiers

• Continuing and inspiring

• A constant impetus to educate and train

Identify, develop, and apply advanced technologies to…

• Enable exploration and discovery

• Allow the public to actively participate in the journey

• Translate the benefits of these technologies to improve life on Earth

Harness the brain power

• Engage the nation’s science and engineering assets

• Motivate successive generations of students to pursue science, math, engineering and technology

• Create the tools to facilitate broad national technical participation

NASA/CP—2005-213655/VOL1 34

-This is a snap shot of the activities that NASA is engaged in across the agency

-Across the Agency everyone is involved and an intricate part of the success for the Vision

*Possible tailoring to the centers as what piece of the plan each center is involved with

This chart shows how comprehensive the Vision is. It encompasses Science, missions – human and robotic – and includes the Exploration Systems Mission Directorate’s portion.

NASA/CP—2005-213655/VOL1 35

Preparing for Mars ExplorationMoon as a test bed to reduce risk for future human Mars

missions• Technology advancement reduces mission costs and

supportsexpanded human exploration

• Systems testing and technology test beds to develop reliability in harsh environments.

• Expand mission and science surface operationsexperience and techniques

• Human and machine collaboration: Machines serve as an extension of human explorers, together achieving more than either can do alone

• Breaking the bonds of dependence on Earth: (e.g./Life Science/Closed loop life support tests)

• Power generation and propulsion development and testing• Common investments in hardware systems for Moon, Mars

and other space objectives

-Traveling to the Moon is important as it enables us to set test beds to go to Mars to minimize risk.

-Establishing an infrastructure by going to the Moon will build sustainability, affordability, and establish credibility.

-Reiterates points outlined

Questions:

Why are we going to the Moon and Mars? Resources. Life Science.

NASA/CP—2005-213655/VOL1 36

Evolutionary AcquisitionINCREMENT 2

INCREMENT 3

• Urgency of requirement

• Maturity of key technologies

• Interoperability, supportability, and affordability of alternative acquisition approaches

• Cost/Benefit of evolutionary vs. single step approach

Key ConsiderationsKey Considerations

Single Step to Single Step to Full Capability ?Full Capability ?

OROR

Evolutionary Acquisition

• Spiral Development: The end-state requirements are not known at program initiation. Those

requirements are refined through system development and demonstration, risk management and continuous user feedback

• Incremental Development: The end-state requirement is known, and that requirement is met over time by developing several increments, each dependent on available mature technology and

resources

Evolutionary Acquisition

• Spiral Development: The end-state requirements are not known at program initiation. Those

requirements are refined through system development and demonstration, risk management and continuous user feedback

• Incremental Development: The end-state requirement is known, and that requirement is met over time by developing several increments, each dependent on available mature technology and

resources

Evolutionary Acquisition allows us to deliver a capability in increments, understanding that there is a need for future capability improvements. It allows us to time phase our requirements and integrate matured technologies into future increments to provide ever increasing capabilities. It is a method that allows for requirement uncertainty due to the inherent phasing of the development.

Evolutionary Acquisition directly contrasts the single step acquisition processes. In the single step acquisition process all of the requirements and the end state must be known up front. This has led to long lead acquisitions and changing requirements that are not able to benefit from maturing technology.

The two approaches utilized in evolutionary acquisition include Spiral Development and Incremental Development. These two approaches need not be used in separation from each other. In our case we are using the two in combination. We are acquiring the systems necessary to go to Mars using spiral development. We do not know the end point of the system capabilities required to go to Mars, but we do know that we need to have crewed flight, which is the first spiral. Within that spiral we have an increment to develop a CEV. We also know we will have a spiral to take us to the moon between 2015 and 2020. The capabilities required for this spiral are still being defined. Once they are defined one or more increments that are managed as unique acquisitions will be used.

NASA/CP—2005-213655/VOL1 37

Strategy-to-Task-to-Technology Process

Required Features &Characteristics

Nation’s Vision

NSPD

Mission Concepts &Requirements

Science Objectives &Concepts of Operations

Tasks & TechnologyRoadmaps

PROGRAM

MISSION

ENGINEERING

DeficienciesDeficiencies

TradeTradeStudiesStudies

Investment Plan

Operational Environments

Available Technologies

PROGRAM

MISSION

OPERATIONS

Modeling& Modeling& Simulation Simulation

Modeling/Simulation

SystemRequirementDocuments

AffordableSystem Design& Development

Specific program tasks derive from management rigor – a disciplined approach to management that includes an acquisition and investment plan targeted toward building new capabilities and engaging in essential research and development. The process of flowing from our strategy to program tasks is iterative. Our strategy for ensuring affordable and sustainable design and development requires extensive modeling and simulation of concepts and their interactions within a range of anticipated operational environments. Cost, performance, and risk data are evaluated iteratively to determine the optimal:

•Requirements set and priorities

•Design for desired capabilities

•Acquisition plan for new capabilities

•Investment plan for research and technology development.

Like our overall efforts, this strategy-to-task process is spiral in nature in that, through repeated cost analysis and performance options, trends and results –including progress in developing specific capabilities and progress in maturing essential technologies – we spiral towards the deployment of new transformational capabilities in a manner that is safe, effective, and affordable.

NASA/CP—2005-213655/VOL1 38

Requirements and TechnologyInvestment Flow

Requirements

Prometheus

Constellation

Technology Maturation

Directorates

Directorates

Prototype Block I Block IISpiral Development:

Advanced Planning

Technologies

MatureDesign

Slide shows the process and flow kicked off by requirements definition. The other directorates and advance planning will provide requirements. This flow helps develop and identify the technology in which we need to invest. Once the technologies are to a point of maturity, they can be integrated.

Analogy: If a car company wants to release a new model, they first build a prototype to be used as their concept car. When the concept is ready for manufacture, it moves to Block 1, which is similar to the first release of the car to the general public – say, a 2005. After gathering data on things that can be added, improved, repaired, or streamlined, the car manufacturer begins work on the next model, most likely to be released in 2006. This continuous improvement helps bring the design closer to maturity in a logical, controlled fashion based upon solid data. In addition, consumers who get a positive feeling from the 2005 model will tell friends and family and the market will be more willing to invest additional funds in a new, improved 2006 model. This allows the manufacturer to add more features and options to their car, just as NASA will be able to add more and more technology to projects as they evolve.

NASA/CP—2005-213655/VOL1 39

Cross-Agency, System of Systems Integration(Lunar Architecture – Illustrative Example Only)

QuickTime™ and aGraphics decompressor

are needed to see this picture.

Transit andLaunch Systems

CrewTransport

Launch

CrewSupport

QuickTime™ and aGraphics decompressor

are needed to see this picture.

Supporting Research

Resource Identificationand Characterization

Biomedical Countermeasuresand Limits

Surface andOrbital Systems

LandingSystems

SurfaceMobility

Comm/Nav

Commonality/EvolvabilityFor Future Missions

MarsCandidates

TelescopeCandidates

Outer Moons Candidates

Technology Options

Long-DurationHabitation

Surface Power andResource Utilization

Pre-PositionedPropellants

Level 0, 1…

Requirements

Requirements

Production &Deployment

Operations& Support

IOCA

SystemIntegration

SystemDemonstration

System Development and DemonstrationDesign

ReadinessReview

FRPDecision

TechnologyDevelopment

OT&E

B C

Concept Refinement

ConceptDecision

ProgramInitiation


Operations& Support

IOCAA

SystemIntegration

SystemDemonstration


ReadinessReview

FRPDecision


OT&E

BB CC

Concept Refinement

ConceptDecision

ProgramInitiation

Level 0, 1É


Operations& Support

IOCA

SystemIntegration

SystemDemonstration


ReadinessReview

FRPDecision


OT&E

B C

Concept Refinement

ConceptDecision

ProgramInitiation


Operations& Support

IOCAA

SystemIntegration

SystemDemonstration


ReadinessReview

FRPDecision


OT&E

BB CC

Concept Refinement

ConceptDecision

ProgramInitiation

Spiral nth?

VisionVision

Level 0, 1É

Requirements

Requirements


Operations& Support

IOCA

SystemIntegration

SystemDemonstration


ReadinessReview

FRPDecision


OT&E

B C

Concept Refinement

ConceptDecision

ProgramInitiation


Operations& Support

IOCAA

SystemIntegration

SystemDemonstration


ReadinessReview

FRPDecision


OT&E

BB CC

Concept Refinement

ConceptDecision

ProgramInitiation

Level 0, 1É


Operations& Support

IOCA

SystemIntegration

SystemDemonstration


ReadinessReview

FRPDecision


OT&E

B C

Concept Refinement

ConceptDecision

ProgramInitiation


Operations& Support

IOCAA

SystemIntegration

SystemDemonstration


ReadinessReview

FRPDecision


OT&E

BB CC

Concept Refinement

ConceptDecision

ProgramInitiation

Spiral nth?Spiral nth?

VisionVision

System of SystemsIntegration

Explain that each of the areas mentioned are subject to their own individual spirals, such as the phases involved in creating the CEV.

This slide exhibits the system of systems approached in ESMD, but for all of the technology and technical prowess, it is missing one key element. (Note: recommend adding picture of human using animation on this slide.) Humans are a key piece of functionality in any NASA system of systems.

NASA/CP—2005-213655/VOL1 40

Constellation Systems (Crew Exploration Vehicle)

CrewedFlight

CrewedFlight

Moon(2015-2020)

Moon(2015-2020)

Spiral 1

Spiral 2

Level 0, 1É

Requirements

Requirements

Non-advocacy ReviewsIndependent Cost Reviews

Spiral nth?

04 05 06 07 08 09 10 11 12 13 14 15 16 17 18 19 2004 05 06 07 08 09 10 11 12 13 14 15 16 17 18 19 20

VisionVision � CEV Demo� 1st Launch Lunar Robotic Orbiter

1st CrewedCEV Flt

1st HumanMoon Mission

1st UncrewedCEV Flt

Mars(2020+)

Mars(2020+)


Operations& Support

IOCA

SystemIntegration

SystemDemonstration


ReadinessReview

FRPDecision


OT&E

B C

Concept Refinement

ConceptDecision

ProgramInitiation


Operations& Support

IOCAA

SystemIntegration

SystemDemonstration


ReadinessReview

FRPDecision


OT&E

BB CC

Concept Refinement

ConceptDecision

ProgramInitiation

Spiral Nth

Systems Engineering

Critical Milestones

Level 0, 1É

System Demonstration

System Integration

DesignReadiness Review

CB

SRR SFR PDR CDR

System Demonstration

System Integration

DesignReadiness Review

CB

SRRSRR SFRSFR PDRPDR CDRCDR


Operations& Support

IOCA

SystemIntegration

SystemDemonstration


ReadinessReview

FRPDecision


OT&E

B C

Concept Refinement

ConceptDecision

ProgramInitiation


Operations& Support

IOCAA

SystemIntegration

SystemDemonstration


ReadinessReview

FRPDecision


OT&E

BB CC

Concept Refinement

ConceptDecision

ProgramInitiation

Shows the systems engineering rigor of the program

Shows overall integration of all spirals

Spiral develop and milestones within each spiral, rigorous implementation plans

How systems engineering permeates throughout each spiral and between spirals

NASA/CP—2005-213655/VOL1 41

Exploration Systems Mission Directorate

Assistant AAAdministration

Deputy AASystems Integration

BusinessOperations

Research &Tech Development

CapabilityDevelopment

Procurement

Deputy AADev

Programs/PEO

Deputy AAResearch

RequirementsFormulation

Program/Risk Assessment

Vacant

Strategic PlanningP. Sunshine

EducationB. McClain

ISS ProgramScientist

D. Thomas

SI ManagerL. Guerra

Sr AdvisorT. Lomax

Deputy PEOM. Hecker

Outreach/Congressional

LiaisonD. Ladwig

Info Mgmt &OperationsC. McCaslin Centennial Challenges

B. Sponberg

Associate Administrator (AA) for

Exploration Systems

Deputy AA

NASA/CP—2005-213655/VOL1 42

Development Programs

for Exploration SystemsDeputy AA

Development Programs/ Program Executive Officer (PEO)

CEV

Space TransSystems

Supporting Surface Systems

SupportingIn-space Systems

Nehman’s AcqBus to Hell

Exploration SystemsResearch & Tech

Human SystemResearch & Tech

Note 1 Transition Programs: X-37, Orbital Express,, DART, PAD, NGLT, OSP

Research & Technology

Development

HubbleServiceMission

ConstellationSystems

TransitionPrograms1

Nuclear Technology & Demonstration

Advanced Systems& Technology

Jupiter Icy Moons Orbiter

Capability Development

Research &Technology Development

Our Development Programs Division develops capabilities and their supporting technologies to sustained and affordable human and robotic exploration. Our Development team is responsible for implementing the requirements and developing the systems to realize the Vision.

This chart demonstrates that the things that we are working in the Research & Technology Development and Nuclear Technology & Demonstration directly feed into Constellation System

NASA/CP—2005-213655/VOL1 43

It is affordable and sustainable

• Paced by experience, technology readiness and flexibility

• Establishing Stepping Stones

• Developing Building Blocks –technology to enable each successive step

• Employing New Approaches – spiral development – build and test

• Fiscal Acquisition Management – Disciplined

It is focused and achievable

• Responds to the nation’s call for a long term space vision

• We have an integrated agency approach

• We have the talent, experience and leadership – recent successes and demonstrated management reforms

• We have the passion and commitment to succeed

-

One Step at a Time

We are implementing the Vision one step at a time

NASA/CP—2005-213655/VOL1 44

Major Challenges

• Integration of piecesBAAs / Trades / RFP / Acq Plan / Budget / International

Detailed Plan - Disciplined Execution

• Refinement of Organization to ExecuteIdentify Needs -- Recruit

• IT Tools to ManageSupport In-House / Buy

• Cross Agency Integration with IndustryStick to Plan -- Communicate

• Maintain Credibility with all Shareholders• Sustainability with Budget Changes

Plan Contingencies

NASA/CP—2005-213655/VOL1 45

Thoughts

• We Have a Great Team – Greater Than the Sum of Its Parts

• We Are Putting Together “World Class” Programs and Processes

• We Are Substantially Changing the Way NASA Does Business by Infusing

Management Rigor, Consistency of Purpose, and Disciplined Processes

• Tremendous Support From Our Administrator, the White House and Growing

Support in Our Congress

• Great Enthusiasm From US and International Industry to Participate

• We Have the Privilege to Be Working on Programs of High National Importance

on Behalf of All Americans

• We Are Inspiring Our Children - the Next Generation of Americans That Will

Pick up the Baton of Exploration

• United Support of Vision will Ensure Sustainability

NASA/CP—2005-213655/VOL1 46

OVERVIEW OF THE ULTRA-EFFICIENT ENGINE TECHNOLOGY AND QUIET AIRCRAFT TECHNOLOGY PROJECTS

Carol Ginty

National Aeronautics and Space Administration Glenn Research Center

Cleveland, Ohio

November 9, 2004

Carol GintyVehicle Systems Projects Office

Acting Deputy ChiefNASA GRC, Cleveland, Ohio

Overview of the Ultra-Efficient Engine Technology

and Quiet Aircraft Technology

Projects

NASA/CP—2005-213655/VOL1 47

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NASA/CP—2005-213655/VOL1 48

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NASA/CP—2005-213655/VOL1 49

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NASA/CP—2005-213655/VOL1 50

Pro

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no

log

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as

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NASA/CP—2005-213655/VOL1 51

Exp

lore

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NASA/CP—2005-213655/VOL1 52

Str

ate

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nvir

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NASA/CP—2005-213655/VOL1 53

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NASA/CP—2005-213655/VOL1 54

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NASA/CP—2005-213655/VOL1 57

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Red

uce

d F

uel B

urn

Tra

nsp

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ing

(RFB

TW)

Sw

ept h

igh-

spe

ed s

lotte

d w

ing

tech

nol

ogy

(TR

L 5

)

Leve

l III

RFB

TW

Aero

dyn

am

ic G

roun

d-to

-Flig

ht S

calin

g G

uid

elin

es C

omp

lete

d

Flig

ht d

emo

nst

ratio

n o

f se

lect

ed (T

BD

) ef

ficie

nt a

ero

dyna

mic

tech

nolo

gy

(TR

L 6

)

Aero

dyn

am

ic g

roun

d-to

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ht s

calin

g g

uid

elil

nes

(TR

L 6

) (in

L2

Pla

n, w

as

not

in T

CA

T/E

AS

I L3

pla

n a

s k

ey

del

ive

Cha

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eriz

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n o

f flig

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yna

mic

s &

co

ntro

l cha

ract

eris

tics

of a

BW

B-c

las

s ve

hic

le (T

RL

5)

Tra

nson

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dap

tive

Bu

mp

tech

no

log

y ( T

RL

5)

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nson

ic W

ing

Ad

aptiv

e B

ump

Te

chn

olo

gy D

em

ons

tratio

n C

om

ple

ted

Sw

ept H

igh

Spe

ed S

lotte

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ing

Te

chn

olo

gy A

sse

ssm

ent C

omp

lete

d

The

pu

rpo

se o

f the

Red

uce

d F

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is to

de

velo

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nd d

em

ons

trate

adv

ance

d a

ero

dyna

mic

a

irfra

me

com

pone

nt te

chn

olo

gie

s a

nd in

tegr

ate

d co

nfig

urat

ion

conc

ept

s e

nab

ling

impr

oved

ae

rody

na

mic

effi

cien

cy fo

r re

duc

ed

fuel

bur

n a

nd/o

r e

nhan

ced

mob

ility

. Te

chn

olo

gie

s w

ill d

irec

tly s

uppo

rt th

e E

ASI p

roje

ct g

oals

of e

nha

ncin

g p

ublic

m

obi

lity

and

en

abl

ing

red

uce

d C

O2/

AS

M e

mis

sio

ns b

y XX

% f

or a

rep

res

enta

tive

20

22 fl

eet r

ela

tive

to a

ba

selin

e 1

997

flee

t. D

eve

lopm

ent a

nd

dem

ons

tratio

n w

ill b

e u

p to

an

d in

clud

ing

TR

L 6

, dep

endi

ng

on

the

tech

nol

ogy

.

The

RF

BTW

Su

bP

roje

ct w

ill d

eve

lop

an

d d

emo

nstra

te a

dvan

ced

ae

rody

nam

ic a

irfra

me

com

pon

ent t

ech

nol

og

ies

an

d in

tegr

ate

d co

nfig

urat

ion

conc

ept

s th

at w

ill e

nla

rge

the

des

ign

sp

ace

for

futu

re v

ehic

les

. Spe

cific

ally

, te

chn

olo

gie

s w

ill ta

rge

t m

eetin

g fu

ture

ve

hicl

e ca

pabi

lity

goa

ls fo

r: 1

) sub

son

ic tr

ansp

orts

with

ext

rem

e cr

uis

e e

ffici

enc

y, 2

) ex

trem

ely

sho

rt ta

ke-

off/l

and

ing

air

cra

ft w

ith re

volu

tion

ary

hig

h s

pee

d cr

uis

e e

ffici

ency

, an

d 3

) hig

h e

ffice

ncy/

low

boo

m s

uper

soni

c cr

uis

e a

ircra

ft.

The

key

con

trib

utio

n o

f the

Sub

Pro

ject

is e

xpe

rime

nta

lly d

em

ons

trate

d te

chno

log

y be

nef

it co

mb

ine

d w

ith d

esi

gn

gui

del

ines

an

d a

n u

nder

sta

ndin

g/r

educ

tion

of m

ulti

dis

cipl

inar

y ri

sk to

pra

ctic

al im

ple

me

nta

tion

. Te

chn

olo

gy tr

ans

fer

will

be

sig

nifi

can

tly

The

RFB

TW

Su

bPro

ject

de

velo

ps

and

de

mon

stra

tes

ad

van

ced

ae

rod

yna

mic

air

fram

e c

om

pon

ent t

ech

nolo

gie

s a

nd

inte

gra

ted

con

figu

ratio

n c

once

pts

ena

blin

g im

pro

ved

aer

odyn

amic

effi

cie

ncy

for r

edu

ced

fue

l bu

rn a

nd

enh

ance

d m

obili

ty o

f fu

ture

airc

raft.

Tec

hno

log

ies

are

sel

ect

ed b

ase

d o

n th

e p

ote

ntia

l im

pact

tow

ard

s e

nabl

ing

the

Veh

icle

Sys

tem

s P

rog

ram

-d

efin

ed fu

ture

ve

hicl

e ca

pabi

lity

goa

ls r

equ

irin

g re

volu

tiona

ry im

prov

eme

nt in

ae

rody

nam

ic e

ffice

ncy

; exa

mp

les

of

tech

nol

ogi

es

incl

ud

e h

igh

sp

eed

slo

tted

win

gs,

tran

son

ic a

dapt

ive

bum

ps,

and

ve

hicl

e co

nce

pts

su

ch a

s th

e b

len

ded-

win

g-

bod

y. S

tate

-of-

the

-art

and

em

ergi

ng

adv

an

ced

co

mp

uta

tion

al to

ols

are

use

d to

de

velo

p th

e va

riou

s te

chn

olo

gie

s, w

hich

are

ve

rifie

d th

rou

gh g

rou

nd-b

ase

d te

sts

up

to fl

igh

t Mac

h a

nd

Rey

nold

s n

um

ber.

Th

e fo

cus

is o

n e

xpe

rime

nta

l de

mo

nstra

tion

of

airc

raft-

enab

ling

tech

nolo

gy

at r

ele

vant

flig

ht c

ond

itio

ns, a

nd

est

abl

ishi

ng

pro

cess

es to

des

ign

an

d a

naly

ze b

ene

fits

of t

he

te

chn

olo

gy; s

ele

cted

tech

nolo

gie

s m

ay b

e d

em

onst

rate

d to

TR

L 6

by

fligh

t tes

t. M

ulti

dis

cip

lina

ry, f

ull-

flig

ht-e

nve

lop

ri

sks

/ba

rrie

rs to

pra

ctic

al/v

iab

le te

chn

olo

gy im

ple

men

tatio

n w

ill b

e id

ent

ifie

d a

nd a

ddr

ess

ed; i

n c

ritic

al r

isk

are

as,

det

aile

d

inve

stig

atio

ns

will

cha

ract

eriz

e a

nd/

or r

edu

ce r

isks

suc

h th

at t

he te

chno

log

y is

und

erst

oo

d a

nd re

adily

tra

nsfe

rred

to in

dust

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AS

trat

egic

Pla

n

Th

e P

resi

den

t’s

Man

agem

ent

Ag

end

a

ES

2020

3.04

:R

educ

e a

ctua

tor

size

to

elim

inat

e a

ctua

tor

“bum

ps.”

ES

2020

3.05

: O

ptim

ize

airf

oil s

ect

ion

& w

ing

desi

gn.

ES

2020

3.06

: In

teg

rate

air

fra

me

& p

rop

ulsi

on

syst

em

to

red

uce

inte

rfe

renc

e d

rag

.

Larg

e-S

cale

, G

roun

d D

emo

CL=

8, L

/D=1

4

0506

07

08

0910

1112

13

141

516

1718

0419

GO

AL

ES

1: I

ncre

ase

Max

imum

Pow

ered

-Lift

Coe

ffici

ent (

CL m

ax

= 10

, SO

A =

7)

Inte

gra

ted

Po

were

d L

ift/

Eff

icie

nt

Cru

ise

Ro

ad

map

Flig

ht D

emo

CL=

8, L

/D=1

4Fl

ight

Dem

oC

L=10

, L/D

=16

Pow

ered

-Lif

tTe

chno

logy

Dow

n-se

lect

GO

AL

ES

2: E

ffic

ient

Cru

ise

(L/D

= 1

6, S

OA

= 1

2)

ES

101:

Incr

ease

circ

ulat

ion

of th

e po

wer

ed

high

-lift

syst

em b

y 50

%.

ES

1010

1.02

:In

vest

igat

e n

ove

l co

nfig

ura

tion

s e

nab

ling

hig

h-li

ft a

nd h

igh-

spee

d cr

uis

e.

ES

1010

1.01

:In

cre

ase

circ

ulat

ion

& r

educ

e/e

limin

ate

se

para

tion

by c

om

bini

ng m

echa

nica

l & f

luid

ic

mea

ns (

mic

ro-a

dap

tive

flow

co

ntro

ls, b

low

ing

, et

c.).

ES

1010

1.03

:In

vest

igat

e t

hru

st v

ecto

ring

/ aug

me

ntat

ion

(US

B, E

BF

, IB

F,

ejec

tors

, flu

idic

, etc

.).

ES

202:

Red

uce

drag

of a

pow

ered

-lift

conf

igur

ed a

ircra

ft by

33%

at a

cr

uise

Mac

h#

of 0

.8.

SS

A

Indu

stry

/Aca

dem

ia

Fun

ded

O

verg

uid

e

Cro

sscu

t U

nfun

ded

Lev

era

ged

NA

SA

OG

A

VS

P F

undi

ng f

or G

oals

ES

1&

ES

2: b

y %

16.5

17.

920

.821

.723

.1 b

y pe

rce

ntag

e fo

r th

e f

irst f

ive

ye

ars

onl

y

$ 47

.3%

$ 10

.5%

$ 27

.9%

$ 10

.5%

$ 3.

8%

Th

e S

yste

ms

Vee

Lin

kag

e:

the W

ho

le P

ictu

re

NASA/CP—2005-213655/VOL1 58

VS

: P

rog

ram

Ou

tco

mes,

FY

05

-09

En

viro

nm

ent:

•70

% N

Ox

Red

uctio

n fo

r su

bson

ic tr

ansp

orts

•P

lus

foun

datio

n te

chno

logi

es fo

r an

add

ition

al 1

0% r

educ

tion

•25

% C

O2

Red

uctio

n fo

r su

bson

ic tr

ansp

orts

•P

lus

foun

datio

n te

chno

logi

es fo

r an

add

ition

al 2

5% r

educ

tion

•10

dB

Noi

se R

educ

tion

for

subs

onic

tran

spor

ts

•P

lus

foun

datio

n te

chno

logi

es fo

r an

add

ition

al 1

0 dB

red

uctio

n

•24

dB

noi

se r

educ

tion

for

adva

nced

gen

eral

avi

atio

n ai

rcra

ft

•D

efin

e “a

ccep

tabl

e so

nic

boom

”fo

r ov

erla

nd s

uper

-cru

isin

g ai

rcra

ft

NASA/CP—2005-213655/VOL1 59

UE

ET

-U

ltra

-Eff

icie

nt

En

gin

e T

ech

no

log

y P

roje

ct

NASA/CP—2005-213655/VOL1 60

UEET O

verv

iew

•S

trate

gic

Fo

cus:

En

vir

on

men

tall

y F

rien

dly

, C

lean

Bu

rnin

g

En

gin

es

–D

evel

op in

nova

tive

tech

nolo

gies

to e

nabl

e in

telli

gent

turb

ine

engi

nes

that

sig

nific

antly

red

uce

harm

ful e

mis

sion

s w

hile

mai

ntai

ning

hig

h pe

rfor

man

ce a

nd in

crea

sing

rel

iabi

lity.

•C

ross

-cu

ttin

g t

ech

no

log

y d

evelo

pm

en

t &

matu

rati

on

fo

cuse

d o

n t

he S

ub

son

ic,

Su

pers

on

ic a

nd

Ro

torc

raft

–S

T,

“R

educ

ed N

Ox

and

Impr

oved

TS

FC

& T

/W”

–S

SA

, “

Hig

h P

erfo

rman

ce In

lets

”–

RC

, “

Adv

ance

d D

rive

Sys

tem

s”

NASA/CP—2005-213655/VOL1 61

Pu

rpo

se(s

)•

Develo

p a

nd

dem

on

stra

te p

rop

uls

ion

tech

no

log

ies

that

ad

dre

ss t

he

go

als

of

the S

ub

son

ic,

Su

pers

on

ic,

an

d R

oto

rcra

ft S

ect

ors

in

th

e V

eh

icle

Syst

em

s P

rog

ram

.

•D

evelo

p p

rop

uls

ion

tech

no

log

ies

to r

ed

uce

Th

rust

Sp

eci

fic

Fu

el C

on

sum

pti

on

(T

SFC

) an

d t

o in

crease

Th

rust

/W

eig

ht

(T/

W)

that

will

resu

lt in

fu

el b

urn

re

du

ctio

ns

of

up

to

15

%,

eq

uati

ng

to

15

% r

ed

uct

ion

s in

CO

2.

•D

evelo

p c

om

bu

sto

r te

chn

olo

gie

s (c

on

fig

ura

tio

n a

nd

mate

rials

) th

at

will

en

ab

le r

ed

uct

ion

s in

Lan

din

g/Take-O

ff (

LTO

) N

Ox

of

70

% r

ela

tive t

o 1

99

6

ICA

O (

Inte

rnati

on

al C

ivil A

via

tio

n O

rgan

izati

on

) st

an

dard

s.

NASA/CP—2005-213655/VOL1 62

UEET

Ult

ra E

ffic

ien

t En

gin

e T

ech

no

log

y

5-y

ear

Pro

ject

Low

Em

issi

ons

Com

bust

or

Hig

hly

Inte

gra

ted I

nle

t

Hig

hly

Load

ed.

Light

Wei

ght

Com

pre

ssor

&

Turb

ines

Inte

lligen

t Pr

opuls

ion

Sys

tem

Foundat

ion

Tec

hnolo

gie

s

Sys

tem

s In

tegra

tion

and D

emonst

ration

Adva

nce

d D

rive

Sys

tem

s fo

r H

eavy

-lift

Roto

rcra

ft

NASA/CP—2005-213655/VOL1 63

Org

an

izati

on

al B

reakd

ow

n S

tru

ctu

re

Ult

ra-E

ffic

ien

t E

ng

ine

Tec

hn

olo

gy

(UE

ET

)M

anag

er:

Kes

tuti

s C

ivin

skas

Act

ing

Dep

uty

: M

ary

Jo L

on

g-D

avis

Hig

hly

Lo

aded

L

igh

t W

eig

ht

Co

mp

ress

ors

an

d

Tu

rbin

esM

anag

er:

Car

ol G

inty

GR

C

Sys

tem

s In

teg

rati

on

&

Dem

on

stra

tio

nM

anag

er:

M

ary

Jo L

on

g-D

avis

GR

C

Inte

llig

ent

Pro

pu

lsio

n S

yste

ms

Fo

un

dat

ion

T

ech

no

log

ies

Man

ager

:

C

aro

lyn

Mer

cer

GR

C

Veh

icle

Sys

tem

s P

rog

ram

(V

SP

)M

anag

er:

Dr.

Ric

har

d W

lezi

en

Veh

icle

Inte

gra

tio

n, S

yste

m &

T

ech

no

log

y A

sses

smen

t (V

IST

A)

Man

ager

: R

ob

ert

McK

inle

y

Bu

sin

ess

Su

pp

ort

Lea

d:

Kar

en C

ran

dal

l

Lo

w E

mis

sio

ns

Co

mb

ust

or

Man

ager

:

Jo

hn

Ro

hd

eG

RC

Hig

hly

Inte

gra

ted

In

let

Man

ager

:M

ich

ael W

atts

LaR

C

Ad

van

ced

Dri

ve

Sys

tem

s fo

r H

eavy

-L

ift

Ro

torc

raft

Man

ager

:

Mar

y Jo

Lo

ng

-Dav

isG

RC

NASA/CP—2005-213655/VOL1 64

Veh

icle

Sect

ors

Su

pp

ort

ed

(1

of

2)

ST

41

42

3.3

2 -

Flo

w m

an

ag

em

en

t an

d co

ntr

ol te

chn

iqu

es

SS

10

20

2.0

3 -

Vari

ab

le f

idelity

in

teg

rate

d M

DO

tools

for

inn

ovati

ve

low

boom

con

fig

ura

tion

s w

ith

hig

h L

/D

Su

bso

nic

(S

T)

Su

pers

on

ic

(SS

)

2.7

.4H

igh

ly

Inte

gra

ted

In

lets

ST

41

22

0.3

0 -

Ad

van

ced

mate

rials

an

d c

oolin

g t

ech

nolo

gie

s fo

r re

du

ced

co

olin

g

ST

41

42

4.3

4 -

Imp

roved

mate

rials

wit

h h

igh

er

tem

pera

ture

ca

pab

ilit

y a

nd

str

en

gth

per

weig

ht

ST

41

22

0.3

1 -

Ad

van

ced

en

gin

e c

om

pon

en

ts

ST

41

22

0.2

9 -

Measu

rem

en

t to

ols

an

d v

alid

ate

d p

hysi

cs-b

ase

d

cod

es

for

hig

hly

load

ed

tu

rbom

ach

inery

SS

51

11

9.2

9 -

Hig

h t

em

pera

ture

lig

ht

weig

ht

mate

rials

, st

ruct

ura

l ass

em

blies

an

d t

herm

al/

en

vir

on

men

tal b

arr

ier

coati

ng

s

SS

71

62

3-3

4 -

Cre

ep

resi

stan

t h

igh

tem

pera

ture

mate

rials

Su

bso

nic

(S

T)

Su

pers

on

ic

(SS

)

2.7

.3H

igh

ly L

oad

ed

Lig

ht

Weig

ht

Co

mp

ress

ors

an

d T

urb

ines

ST

61

93

2.4

1 -

Ad

van

ced

co

mb

ust

or

tech

niq

ues

to r

ed

uce

NO

x

ST

61

93

2.4

3 -

Ad

van

ced

mate

rials

fo

r re

du

ced

co

mb

ust

ion

lin

er

coolin

g

ST

61

93

2.4

5 -

Valid

ate

d c

om

bu

stio

n c

od

es

for

desi

gn

an

d c

on

trol

of

low

em

issi

on

s co

mb

ust

ors

Su

bso

nic

(S

T)

2.7

.2

Lo

w E

mis

sio

ns

Co

mb

ust

or

Ap

pro

ach

es

Sect

or

WB

S

NASA/CP—2005-213655/VOL1 65

Veh

icle

Sect

ors

Su

pp

ort

ed

(2

of

2)

ST

41

22

0.2

9 -

Measu

rem

en

t to

ols

an

d v

alid

ate

d p

hysi

cs-b

ase

d c

od

es

for

hig

hly

-lo

ad

ed

tu

rbo

mach

inery

ST

61

72

7.3

8 -

Measu

rem

en

t to

ols

an

d v

ali

date

d g

lob

al an

d l

oca

l m

od

els

to

im

pro

ve c

ycl

e d

esi

gn

s

ST

41

42

6.3

6 -

Hig

hly

reliab

le lig

ht

weig

ht

mech

an

ical sy

stem

s

SS

51

01

5.2

3 -

Inle

t d

isto

rtio

n a

nd

fan

dyn

am

ic c

on

tro

l

SS

51

11

6.2

4 -

Inle

t o

pera

bilit

y a

nd

vis

cou

s lo

ss r

ed

uct

ion

Su

bso

nic

(S

T)

Su

pers

on

ic

(SS

)

2.7

.5S

yst

em

In

teg

rati

on

an

d

Dem

on

stra

tio

n

TB

D-

Sp

eci

fic

ap

pro

ach

es

will

be i

den

tifi

ed

pen

din

g t

he r

esu

lts

of

aco

mp

eti

tive p

rocu

rem

en

t p

roce

ssS

ub

son

ic

(ST)

Su

pers

on

ic

(SS

)

2.7

.7In

tellig

en

t P

rop

uls

ion

S

yst

em

s Fo

un

dati

on

te

chn

olo

gy

RC

60

91

1.2

6 -

Meth

od

s to

au

tom

ati

cally d

ete

ct f

ailu

re o

f p

ow

er

train

Ro

torc

raft

(R

C)

2.7

.6

Ad

van

ced

Dri

ve

Syst

em

s fo

r H

eavy-l

ift

Ro

torc

raft

Ap

pro

ach

es

Sect

or

WB

S

NASA/CP—2005-213655/VOL1 66

UEET T

ech

no

log

ies

LO

W N

Ox

CO

MB

US

TO

R

TE

CH

NO

LO

GIE

S-

Lea

n B

urn

Inje

ctio

n-

Ric

h B

urn

Inje

ctio

n-

Co

atin

gs

for

met

allic

lin

ers

HIG

H-L

OA

DE

D A

XIA

L

CO

MP

RE

SS

OR

-3D

aer

o-

Inve

rse

des

ign

-R

ecir

cula

tin

g c

asin

g t

reat

men

t

AC

TIV

E F

LO

W C

ON

TR

OL

F

OR

BW

B IN

LE

TS

-S

ynth

etic

jets

HIG

H-W

OR

K S

ING

LE

-ST

AG

E

HIG

H-P

RE

SS

UR

E T

UR

BIN

E-

Lo

w s

ho

ck lo

ss d

esig

n-

Red

uce

d c

oo

ling

loss

es

HIG

H-L

OA

DE

D L

OW

PR

ES

SU

RE

T

UR

BIN

E (

Reg

ion

al)

-H

ub

en

dw

all b

leed

-L

P s

tag

e re

du

ctio

n

PH

YS

ICS

-BA

SE

D M

OD

EL

ING

& C

FD

-A

P20

00 (

mu

ltib

lock

AP

NA

SA

)-

Nat

ion

al C

om

bu

stio

n C

od

e-

MS

U-T

UR

BO

-G

len

n-H

T

UE

ET

Tec

hn

olo

gy

Po

rtfo

lio D

eriv

ed t

o S

up

po

rt V

ehic

le S

yste

ms

Pro

gra

m

Go

als

for

Imp

rove

d T

SF

C a

nd

T/W

an

d R

edu

ced

NO

x

HIG

H-L

OA

DE

D H

P/L

P

TU

RB

INE

SY

ST

EM

-R

edu

ced

inte

ract

ion

loss

-S

yste

m w

eig

ht

red

uct

ion

AD

VA

NC

ED

MA

TE

RIA

LS

-C

eram

ic M

atri

x C

om

po

site

Van

es-

Lo

w C

on

du

ctiv

ity

TB

C-

Po

lym

er M

atri

x C

om

po

site

C

om

po

nen

ts-

Ad

van

ced

Allo

ys

ME

CH

AN

ICA

L C

OM

PO

NE

NT

S-

Ad

van

ced

No

n-C

on

tact

ing

Sea

ls-

Lig

htw

eig

ht

Gea

rbo

x C

om

po

nen

ts

for

Lar

ge

Gea

red

Tu

rbo

fan

En

gin

e

TU

RB

INE

TIP

CL

EA

RA

NC

E S

EN

SO

R-

Mic

row

ave

sen

sor

-C

on

tro

l gap

s to

imp

rove

eff

icie

ncy

NASA/CP—2005-213655/VOL1 67

UEET B

ud

get

Sp

lit

By S

ect

or

Subso

nic

Super

sonic

Roto

rcra

ft

Bu

dg

et:

% b

y S

ect

or

3%R

otor

craf

t

11%

Sup

erso

nic

86%

Sub

soni

c

To

tal d

oes

no

t in

clu

de

the

UE

ET

Pro

ject

’s 2

0% f

un

din

g

of

fou

nd

atio

nal

tec

hn

olo

gie

s.

NASA/CP—2005-213655/VOL1 68

UE

ET

L2

Pla

nS

ecto

r w

ith

Ob

ject

ives

0506

0708

Fis

cal Y

ear

UE

ET

Tec

hA

sses

smen

t

Fin

aliz

e IP

SIn

vest

men

tP

ortfo

lio

Che

ck-o

ut o

fS

ingl

e S

pool

Tur

bine

Fac

ility

Cro

ss C

utt

ing

Su

bso

nic

Tra

nsp

ort

sR

educ

e E

ngin

e N

Ox

emis

sio

ns

ove

r th

e L

TO

cyc

le

Incr

ease

Th

erm

al E

ffic

ien

cy

Red

uce

Bar

e E

ngin

e W

eigh

t

Red

uce

TS

FC

Red

uce

Inst

alla

tion

Pen

altie

s

Red

uce

Par

ticu

late

s an

d A

ero

sols

by

10%

Red

uce

Pod

/Eng

ine

Sub

syst

em W

eigh

t -- 1

5%

Su

per

son

ic A

ircr

aft

Dec

reas

e F

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NASA/CP—2005-213655/VOL1 87

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NASA/CP—2005-213655/VOL1 88

2004 NASA Seal/Secondary Air System Workshop2004 NASA Seal/Secondary Air System Workshop

High Misalignment Carbon Seals

High Misalignment Carbon SealsFor The

Fan Drive Gear System Technologies

Dennis Shaughnessy / Lou Dobek

United Technologies - Pratt & Whitney

East Hartford, Connecticut

860-557-1675 / 860-565-3034

[email protected] / [email protected]

HIGH MISALIGNMENT CARBON SEALS FOR THE FAN DRIVE GEAR SYSTEM TECHNOLOGIES

Dennis Shaughnessy and Lou Dobek

United Technologies Pratt & Whitney

East Hartford, Connecticut

The Ultra Efficient Engine Technology (UEET) program is a NASA supported program to develop and

demonstrate technology for quiet, fuel-efficient, low-emissions next generation commercial gas turbine aircraft engines. An essential role for achieving lower noise levels and higher fuel efficiencies is played by the power transmission gear system connected to the fan. Reduction geared systems driving the fan will be subjected to inertia and gyroscopic forces resulting in extremely high angular and radial misalignments. Because of the high misalignment levels, compartment seals capable of accommodating angularities and eccentricities are required. Pratt & Whitney and Stein Seal Company selected the segmented circumferential carbon seal as the best candidate seal type to operate at highly misaligned conditions and developed a test program to determine misalignment limits of current segmented circumferential seals. The long-term goal is to determine a seal design able to withstand the required misalignment levels and provide design guidelines. A technical approach is presented, including design modification to a “baseline” seal, carbon grade selection, test rig configuration, test plan and data acquisition. Near term research plans are also presented.

NASA/CP—2005-213655/VOL1 89



Background

Tomorrow’s Engines with Geared Fans will be subjected to extreme conditions such as:

• High angular and radial seal misalignmentsGyroscopic loads - angular misalignment

Sun input gear orbiting - radial/eccentric misalignment

• Higher LPC shaft speed; ~10,000 RPM

• Large Diameter Fan Hub

Seals capable of accommodating high misalignment, high rubbing speeds,

low pressure differentials and large diameters must be developed.

Background information on principal causes of extreme conditions in Advanced Commercial Engines. Such conditions impose on seals high misalignment, high rubbing speed, large diameters and low pressure differentials.

NASA/CP—2005-213655/VOL1 90



Program Objective

The use of the reduction gear system as the platform for UEET demonstration engines will provide revolutionary improvements inengine performance, weight, size, and noise. Due to high periodic radial and angular misalignments introduced into the gear system, high misalignment seals are required to provide adequate compartment sealing beyond present capability. These seals must also have adequate life.

Current Phase Objective

Using data and lessons learned from previous test phases, continue testing the advanced seal design to 0.105” with both the baseline material and an alternate material.

Overall program objective identifies the need for seals capable of periodic high radial and angular misalignment.

The current phase objective is to test an advanced seal design up to 0.105” of total misalignment using both the baseline material as well as an alternate material.

NASA/CP—2005-213655/VOL1 91



Gear system must be flexibly connected between low spool and fan shaft

Geared Turbofan Engine (GTF)

Geared Turbo Fan Provides• 3%-4% TSFC improvement over conventional turbofan engines.• 30db noise reduction.

Misalignment seals are located along the flexible shaft between the low spool and fan shaft.

NASA/CP—2005-213655/VOL1 92



FWD. AIR/OILSEAL

REAR AIR/OILSEAL

LPC COMPART-MENT SEAL

Advanced Engine Seal Locations

Angular Misalignment

Eccentricity

High Angular MisalignmentHigh Eccentricity/Oil Flooding

High Speed / Wear Life

Large Diameter Low Pressure Seal

Seal locations within the forward compartments of the fan drive geared engine. Forward air/oil seal represents the location of the highest source of angular and radial misalignment.

NASA/CP—2005-213655/VOL1 93



Seal Operating Conditions

CURRENT FOCUS

FWD. REAR FDGS/LPCAIR/OIL SEAL AIR/OIL SEAL COMPARTMENT SEAL

Required Life (hours) 30,000 30,000 30,000

Delta P (psi) <50 <50 40-50

Surface Speed (ft/s) 33 90 345

Buffer Air Temperature (deg. F) 350 350 415

Angular Misalignment (deg) 0.5 0.2 0.1

Eccentricity (inches) 0.005 0.02 0.005

Sealing Diameter (inches) 2.95 2.95 11.2

Type Segmented/ Segmented/ Segmented/bellows/ other ring/other other

Seal operating conditions (required life, pressure differentials, speeds, misalignment levels and others).

Critical requirements are highlighted.

NASA/CP—2005-213655/VOL1 94



LPT Input Shaft RotationCCW from rear

A

Detail A

FWD Air/Oil SealHigh Misalignment Seal

LPC Compartment Seal

High Misalignment Seal

Sun Gear

Fan Drive Gear System must withstand periodic misalignments as high as 0.105” due to “g” and gyro loads.

Rear Air/Oil Seal

Fan drive gear systems must withstand periodic misalignments as high as 0.105”due to “g” and gyro loads.

NASA/CP—2005-213655/VOL1 95



Approach

Misalignment Seal Test Rig Program

Stein Seal selected as the seal supplier/tester.

Testing at supplier’s facilities.Step 1 – Previous Update• “Baseline” seal design• Carbon grade “X” - high strength, low modulus.• Misalignment increased in steps up to 0.020 in. radial & 0.5° angularStep 2 – Previous Update• Misalignment increased in steps up to 0.040 in. radial & 0.5° angularStep 3• Alternate seal with Carbon grade “X” tested• Misalignment increased in steps up to 0.105 in. radial & 0.5° angular• Alternate seal with Carbon grade “Y” tested• Misalignment increased in steps up to 0.105 in. radial & 0.5° angular

Technical approach of misalignment seal development program. Three main steps will be followed starting from “baseline” seal testing.

NASA/CP—2005-213655/VOL1 96



High Misalignment Test Rig Facility

High Misalignment Test Rig

NASA/CP—2005-213655/VOL1 97



Air Side

Oil Side

Rig Bearing

Drive Section

CL

Test Section

Test SealHeater

Shim

Pilot RingShim

Shims - angular misalignmentPilot ring - radial misalignment

A

Detail A

Radially Eccentric Runner

Runner

Offset

Seal Rig Imposes Radial and Angular Misalignment

Seal test rig schematics used to impose radial and angular misalignment. Shims are used to impose angular misalignment and pilot rings are used to impose radial misalignment.

NASA/CP—2005-213655/VOL1 98



Carbon Seal Designs

1-Piece Segmented Circumferential

(Baseline Design)


(Advanced Design)

Circumferential Seal Ring

Seal Housing

Backplate

Retaining Ring

Garter Spring

Compression Spring

Back Ring Segment

Seal Housing

Backplate

Retaining Plate

Garter Springs

Compression Spring

Seal Ring

Retaining Ring

Cover Ring

• 1 Piece design resulted in wear rates that exceeded project limit goals.

• Base material (Carbon X) is suspect to be inappropriate for this application.

• Alternate design & material need to be investigated.

Baseline seal is compose of a one-piece 4 segmented seal. Alternate design is composed of a three-piece design, each piece consisting of four segments.

NASA/CP—2005-213655/VOL1 99



0.00000

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0.030 0.0

0.060 0.0

0.090 0.0

0.096 0.0

0.096 0.2

0.096 0.2

0.096 0.3

0.096 0.4

0.096 0.5

Misalignment Tests

100

Ho

ur

Wea

r (I

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es)

Back Ring Seal Ring

Carbon X - Radial Wear Results

Carbon X - Axial Wear Results

0.00000

0.00100

0.00200

0.00300

0.00400

0.00500

0.0000.0

0.0300.0

0.0600.0

0.0900.0

0.0960.0

0.0960.2

0.0960.2

0.0960.3

0.0960.4

0.0960.5

Misalignment Test

100

Ho

ur

Wea

r (I

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Back Ring Cover Ring Seal Ring

• Carbon X repeatedly exceeds 100 hour wear limit.

1. Excessive coking indicated negative wear.

2. Test terminated after two attempts resulted in broken seal segments.


Back Ring Segment

Seal Ring

Cover Ring

100 Hour Wear Limit

100 Hour Wear Limit

1 2

2

Carbon X repeatedly exceeds the 100 hour wear limit goal and testing was terminated after multiple failures.

NASA/CP—2005-213655/VOL1 100



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0.00200

0.00300

0.00400

0.00500

0.000 0.0

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0.060 0.0

0.090 0.0

0.096 0.0

0.096 0.2

0.096 0.2

0.096 0.3

0.096 0.4

0.096 0.5

Misalignment Tests

100

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(Inch

es)

Back Ring Seal Ring Carbon Y - Axial Wear Results

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0.00200

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0.0600.0

0.0900.0

0.0960.0

0.0960.2

0.0960.2

0.0960.3

0.0960.4

0.0960.5

Misalignment Test

100

Ho

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Wea

r (I

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Back Ring Cover Ring Seal Ring

Carbon Y - Radial Wear Results

• All components using Carbon Y meets the 100 hour wear limit in Radial misalignment tests.

• Carbon Y components exceed 100 hour wear limit in axial misalignment tests. Mounting hardware damage identified.


Back Ring Segment

Seal Ring

Cover Ring

100 Hour Wear Limit

100 Hour Wear Limit

Carbon Y meets the 100 hour wear limit under purely radial misalignment conditions. Seal retaining hardware suffered fatigue and failure during combination radial & angular misalignment tests. These failures are to be investigated at the potential reasons for the 100 hour wear limit to be exceeded.

NASA/CP—2005-213655/VOL1 101



Backplate tang steadily wearing into slot of Retaining Ring. Current Seal Housing prohibited design change to increase Retaining Ring to the full thickness of the tang.

Back Ring Segment

Seal Housing

Backplate

Retaining Plate

Garter Springs

Compression Spring

Seal Ring

Retaining Ring

Cover Ring

Retaining Ring

Backplate

Photos of seal retaining hardware show the wear the occurred during misalignment testing.

NASA/CP—2005-213655/VOL1 102



Back Ring Segment

Seal Housing

Backplate

Retaining Plate

Garter Springs

Compression Spring

Seal Ring

Retaining Ring

Cover Ring

Backplate failure during test phase of advanced design misalignment tests

Seal Ring wear during test phase of advanced design misalignment tests

Backplate Key steadily wore into slot of Back Ring and Seal Ring. Modification increased the number of Keys from one to two.

Backplate stress crack identified after test with the largest combination of Radial and Angular misalignment. Modification increased the corner radii at the keys.

Photos of seal and associated retaining hardware show the wear and stress fracture that occurred during misalignment testing.

NASA/CP—2005-213655/VOL1 103



Conclusions• The baseline design does not meet wear requirements based on Phase II test

results and should not be further developed.• The optimized three-piece carbon design is a significant improvement over the

baseline seal.• The Carbon Y material appears to offer more consistent results and improved

wear performance than the baseline Carbon X material.• Seal retaining hardware on the 3-piece design worn in several instances and

may explain carbon wear rates that were greater than goal.

Recommendations1. Seal retaining hardware should be scrutinized and optimized for the three-

piece design.2. Misalignment tests should be run on the improved seal retaining hardware

with the three-piece carbon seal, fabricated from Carbon Y, and compared to current test results.

3. Upon successful completion of the misalignment tests, durability testing should be run on the three-piece carbon seals with improved hardware.

Conclusions identify that baseline seal design should not be further developed. Also the 3-piece design is a significant improvement over the baseline design. Carbon Y material appears to offer improved wear results from that of Carbon X. Further work is needed to improved the seal retaining hardware.

Recommendations are to investigate seal retaining hardware and improve. Another battery of tests utilizing Carbon Y and improved retaining hardware should be run and compared to the current results.

NASA/CP—2005-213655/VOL1 104



Plans for Next Year & Beyond

Optimize seal designsFabricate test sealsRadial & angular misalignment testsOptimized seal durability tests

Oil windback designUEET demo engine hardware

Oil windback tests

2005

2006

2007

Plans for continuation include design optimization, durability testing, and windback design and testing.

NASA/CP—2005-213655/VOL1 105



Goal: Develop durable seals with 0.105” misalignment capability.

Schedule:0.040 Baseline

Misalignment Tests

Optimize Seal Design

Design & Mfg Improved Seal

0.105 Misalignment Tests on Baseline & Improved Design

Analysis & Report

Assess Current Cir. Seal Capability - Mfg. Seals for 0.040 Misalignment Tests

Fabricate Test Seals

Radial & Angular Misalignment Tests Durability Tests

UEET Demo Engine Hardware

Oil Windback Development Tests

Oil Windback Design for High Misalignment Seals

2003

2005

2006

2007

Objective – Development of seals capable of 0.105” misalignment for use in the Pratt & Whitney reduction gear system.

NASA/CP—2005-213655/VOL1 106



Base Seal - Radial Misalignment Tests

0.00000

0.00100

0.00200

0.00300

0.00400

0.00500

0.000 0.0

0.000 0.0

0.010 0.0

0.015 0.0

0.020 0.0

0.030 0.0

0.040 0.0

0.040 0.0

Misalignment Test

100

Ho

ur

Wea

r (I

nch

es)

Axial Wear Radial Wear

RadialAngular

Step one and two test results indicate wear rates in excess of those desired. Unexpected reduction in wear rates at higher levels of misalignment under investigation.

NASA/CP—2005-213655/VOL1 107



Base Seal - Angular Misalignment Tests

0.00000

0.00100

0.00200

0.00300

0.00400

0.00500

0.000 0.1

0.000 0.2

0.000 0.3

0.000 0.4

0.000 0.5

0.030 0.5

0.030 0.5

Misalignment Test

100

Ho

ur

Wea

r (I

nch

es)

Axial Wear Radial Wear

RadialAngular

Step one and two test results indicate wear rates in excess of those desired. Unexpected reduction in wear rates at higher levels of misalignment under investigation.

NASA/CP—2005-213655/VOL1 108



Carbon Seal Wear Damage at 0.040” Misalignment

Radial Pad

Radial Lip

Axial Lip

Axial Pad

Anti-rotation slot

Baseline Design Seals

Step one Carbons indicate excessive wear and chipping along the tongue and sockets of the segments.

NASA/CP—2005-213655/VOL1 109



Commercialization Aspects

Fan Drive Gear System development identified certain technologies as key requirements of which, High Misalignment Seals are extremely important.

Geared Turbo Fan Provides:

• 3%-4% TSFC improvement over conventional turbofan engines.

• 30db noise reduction.

Gear System technology also lends itself to Rotorcraft transmissions.

The circumferential seals are Stein Seal designs that are being optimized in this program.

The Fan Drive Gear System offers significant advances in the areas of weight reduction, noise reduction and fuel consumption.

NASA/CP—2005-213655/VOL1 110

LEAKAGE AND POWER LOSS TEST RESULTS FOR COMPETING TURBINE ENGINE SEALS

Margaret P. Proctor


Cleveland, Ohio

Irebert R. Delgado U.S. Army Research Laboratory


Advanced brush and finger seal technologies offer reduced leakage rates over conventional labyrinth seals used

in gas turbine engines. To address engine manufacturers’ concerns about the heat generation and power loss from these contacting seals, brush, finger, and labyrinth seals were tested in the NASA High Speed, High Temperature Turbine Seal Test Rig. Leakage and power loss test results are compared for these competing seals for operating conditions up to 922 K (1200 °F) inlet air temperature, 517 KPa (75 psid) across the seal, and surface velocities up to 366 m/s (1200 ft/s).

Leakage and Power Loss Test Resultsfor Competing Turbine Engine Seals

Margaret P. ProctorGlenn Research Center,

Cleveland, Ohio

Irebert R. DelgadoU. S. Army Research Laboratory,

Glenn Research Center,Cleveland, Ohio

NASA Seal/Secondary Air Delivery WorkshopNASA Glenn Research Center

Cleveland, OhioNovember 9-10, 2004

CD-04-82658

NASA/CP—2005-213655/VOL1 111

Leakage and Power Loss Test Resultsfor Competing Turbine Engine Seals

Margaret P. ProctorGlenn Research Center,

Cleveland, Ohio

Irebert R. DelgadoU. S. Army Research Laboratory,

Glenn Research Center,Cleveland, Ohio

NASA Seal/Secondary Air Delivery WorkshopNASA Glenn Research Center

Cleveland, OhioNovember 9-10, 2004

CD-04-82658

The authors acknowledge contributions of:

Arun Kumar – Development of finger seal tested.

NASA GRC – conducted all testing.

Bruce Steinetz – guided the design, procurement, and fabrication of the High Temperature, High Speed Turbine Seal Test Rig

For the complete story please see “Leakage and Power Loss Test Results for Competing Turbine Engine Seals” paper by M. P. Proctor and I. R. Delgado presented at Turbo Expo 2004 sponsored by the American Society of Mechanical Engineers in Vienna, Austria, June 14-17, 2004, GT2004-53935, NASA/TM-2004-213049, ARL-TR-3157.

NASA/CP—2005-213655/VOL1 112

NASA Glenn Research Center CD-04-82658

Reducing Secondary Air Leakage in Jet Engines

Benefits:• Higher engine performance

– Decreased specific fuel consumption– Increased thrust

• Better investment towards performance gain than components, such ascompressors and turbines.

Some Considerations: Heat Generation and Power Loss• Changes in engine air temperatures from stage to stage cannegatively affect engine efficiencies.

• Friction from contacting seals increases the amount of torque needed.

• Advanced engines operate at very high temperatures. Excessive heatgeneration at the seal could expose downstream components totemperatures that exceed material capabilities.

Self-explanatory

NASA/CP—2005-213655/VOL1 113


Test Results Compared for…

Three Seals:

• 4-Knife Labyrinth

• Brush seal

• Finger seal

Test Rotors:

• 215.9 mm (8.5 in)

• Grainex Mar-M-247

• CrC (HVOF) coating on o.d.

Initial radial clearance Axial length

229.0 μm ( .009 in) 11.20 mm (.44 in)

–96.5 μm (-.0038 in) 4.27 mm (.168 in)

–165.0 μm (-.0065 in) —

Leakage rate, power loss and wear results are compared for 3 seals: 4-knife labyrinth seal, a brush seal, and a finger seal.

The labyrinth seal had an initial radial clearance of 229 microns.

Both the brush seal and the finger seal initially had an interference with the rotor.

The radial interference of the brush seal was 96.5 microns.

The radial interference of the finger seal was 165 microns.

The finger seal has an axial length similar to the brush seal.

The axial length of the brush seal is 38 percent of the labyrinth seal.

The 216 mm test rotors used are made of Grainex Mar M-247 and have a chrome carbide coating on the o.d.

NASA/CP—2005-213655/VOL1 114


Four-Knife Labyrinth Seal

Labyrinth Seal Design Parameters

229 μm

4 knife labyrinth seal has straight thru knife-edges and a radial clearance of 229 microns.

It is made of Inconel 625.

Only static tests were conducted.

It was determined that for safe operation at maximum speed and temperature that the labyrinth seal would need radial clearance of 305 microns at build.

KTK, an labyrinth seal analysis code, was used to predict the flowrates of this larger clearance seal.

NASA/CP—2005-213655/VOL1 115


Brush Seal With Flow Deflector

• Inconel-625 sideplates

• Bristles are:– Haynes-25– 102 μm diameter– At 50° angle to radius– 675 bristles/cm of circumference

at seal i.d.

• Fence height: 1.27 mm

• Total axial thickness: 4.27 mm

This is the brush seal we tested.

A brush seal is simply an annular pack of bristles sandwiched between two annular plates. This brush seal has a flow deflector on the upstream side to prevent the flow from jetting thru the bristles.

The sideplates are made of Inconel 625 and the bristles are made of Haynes-25.

The 102 micron (.004 in) diameter wire bristles are at a 50 degree angle to the radius allowing them to deflect away from the rotor as cantilever beams to accommodate rotor growth or deflections.

There are 675 bristles/cm of circumference at the seal id and the seal has a fence height of 1.27 mm (.050 in).

Fence height is the distance from the bristle i.d. to the i.d. of the seal backplate.

Total axial thickness is 4.27 mm (.168 in).

NASA/CP—2005-213655/VOL1 116


Finger Seal Design

The finger seal functions similarly to the brush seal.

It is made of AMS 5537, a cobalt-base alloy.

It is comprised of forward and aft coverplates, 2 spacers, and 3 finger elements.

The annular finger elements are made of sheet stock and have a series of cuts along the inner diameter to create the fingers.

These fingers deflect as cantilever beams when loaded by rotor interference due to design, thermal or centrifugal growth or rotordynamics.

NASA/CP—2005-213655/VOL1 117


High-Temperature, High-Speed Turbine Seal Rig

All the seals were tested in the High Temperature, High Speed Turbine Seal Rig at NASA Glenn Research Center.

Hot air enters the bottom of the test section, passes thru the seal and exits thru the 2 exhaust lines.

Under some conditions it is necessary to bypass the seal to maintain the test temperature.

Seal supply and bypass flow rates are measured. The difference of these 2 flows is the seal leakage rate.

An air turbine is used to power the rig.

The torquemeter is located between the turbine and the test rig.

A balance piston is used to control the thrust loads on the bearings due to the pressure differential acting on the seal test rotor.

NASA/CP—2005-213655/VOL1 118


Test Seal Configuration and Location of Research Measurements

Here is the (Point out) the Test Rotor, Seal holder, and test seal.

The test seal is held in place by the spacer and the seal clamp.

For some tests proximity probes were mounted in the spacer to measure changes in clearance.

A metal c-seal is used to prevent flow from bypassing the test seal.

Inlet and exit pressures and temperatures are measured as well as the seal backface temperature.

NASA/CP—2005-213655/VOL1 119


Flow Factor

m = air leakage flow rate, kg/s.

Tavg = average seal air inlet temperature, K.

Pu = air pressure upstream of seal, MPa.

Dseal = outside diameter of the test rotor, m.

•

We present the leakage performance of the seals in terms of flow factor. Use of flow factor allows comparison of data taken at different test conditions and between seals of different diameters.

Flow factor is a function of:

the mass leakage rate

average seal air inlet temperature and upstream pressure

and seal diameter

NASA/CP—2005-213655/VOL1 120


Tests Conducted

• 4 hour endurance test: inspections after 1, 2, and 4 hours

This is a table showing the tests conducted.

4 tests were conducted on both the brush seal and finger seal: a static test, a performance test, an endurance test, followed by a post-endurance performance test. For the brush seal, the static test was repeated too.

Data was taken at five temperature conditions up to 922 K.

For the static tests, at each temperature the pressure differential was increased to 517 kPa (75 psid) and then decreased to 0.

In the performance tests, at each temperature, and at each of three pressure levels of 69, 276, and 517 kPa, the speed was stepped up to 366 m/s (1200 ft/s) and the stepped back down.

The 4 hour endurance test was conducted at the maximum conditions of 922 K, 517 kPa, and 366 m/s with inspections after hours 1, 2, and 4.

Only a room temperature static test was conducted for the labyrinth seal.

NASA/CP—2005-213655/VOL1 121


Initial Static Leakage Performance of Brush and Labyrinth Sealswith KTK Predictions

297 K Average Seal Inlet Air Temperature

229 μm

This is a comparison of the initial static leakage performance data measured for the brush seals and the labyrinth seal at room temperature. The plot shows flow factor versus pressure drop across the seal.

Brush seal leakage is 24% less than the labyrinth seal and uses only 38% of the axial space the labyrinth seal requires.

The KTK predicted leakage rate for the seal tested is shown by the solid line.

The measured labyrinth seal leakage is 91% of KTK predicted flow factor of 17.6 kg–√K/MPa-m-s.

For safe operation to 922 K and 366 m/s a labyrinth seal with a 305 μm radial clearance is required.

At that clearance, the KTK predicted flow factor is 25 which is twice the brush seal leakage.

NASA/CP—2005-213655/VOL1 122


Time History of Finger and Brush Seal Endurance Tests

922 K, 366 m/s, 517 kPa

This is a time history of the leakage performance of the brush and finger seals during the endurance test conducted at 922 K, 366 m/s and 517 kPa.

The finger seal flow factor shown by the solid symbols is less than the brush seal through out the test.

Initially the brush seal leakage is only slightly higher than the finger seal.

However, after 4 hours of testing the finger seal leakage flow factor was 36% less than the brush seal flow factor.--------------• The predicted flow factor for 305 μm radial clearance labyrinth seal at these

conditions is 26.1 kg–√K/MPa-m-s, which is 2.16 × brush seal, 3.38 × finger seal.

NASA/CP—2005-213655/VOL1 123


Post-endurance Performance Tests – Finger and Brush SealLeakage Data


After the endurance test the performance test was repeated. This is a comparison of the brush and finger seal leakage performance data at 922 K.

The brush seal flow factor data shown in black is higher than the finger seal data in red.

The finger seal shows a definite pressure closing effect as the flow factor is lower at higher pressure differentials.

At 922 K, 366 m/s, 517 kPa, the flow factors are greater than in the initial performance test by a factor of 1.6 for the Finger Seal and 2.5 for the Brush Seal.

NASA/CP—2005-213655/VOL1 124


Post-Endurance Performance Tests – Finger Seal (FS) and Brush Seal (BS)Average Power Loss Versus Average Surface Speed


Seal power loss was calculated by measuring the torque with a seal and subtracting the tare torque measured without a seal to determine the seal torque and then multiplying the seal torque by the speed.

This is a comparison of the finger and brush seal power loss at 922 K (1200 ºF). The data shown here is from the post-endurance performance test, which is very similar to the first performance test.

As expected the power loss increases with speed and with pressure differential across the seal.

At all three pressure conditions the brush seal power loss is slightly higher than the finger seal power loss.

At the maximum speed and pressure of 366 m/s and 517 kPa the power loss for the brush seal is 10.5 kW and for the finger seal it is 9.72 kW.

We recognize that these seal power loss values may be substantially higher than the true seal power loss since the tare torque was measured at ambient pressure.

NASA/CP—2005-213655/VOL1 125


Finger Seal Approximated True Power Loss Versus Average SurfaceSpeed for Performance Test


Windage losses on the high pressure side of the test rotor and balance piston were approximated using a solution provided by Schlicting. The additional torque on the bearings due to the thrust loads was also estimated. These estimations were then used to determine the approximated true power loss for the finger seal shown here.

The approximated true power loss increases with speed and with pressure, but not to the same degree as the data based on the ambient pressure tare torque.

At the maximum conditions of 922 K, 366 m/s, and 517 kPa the approximated true seal power loss is 2.36 kW. This is 1/4 of the seal power loss based on the ambient pressure tare torque.

Hence, this finding suggests that additional testing with a straight cylindrical seal or labyrinth seal is needed to establish tare torque data that accounts for pressure differentials across the test seal.

Using the ambient pressure tare torque is adequate for comparing performance, however.

NASA/CP—2005-213655/VOL1 126


Finger and Brush Seal Accumulative Weight Loss Versus AccumulativeRun Time

This is a comparison of the finger and brush seal accumulative weight loss versus accumulative run time.

The finger seal had a final weight loss of 6 g and the brush seal lost ~1g.

70–71% of the seal weight loss occurred in the initial performance tests.

Weight loss was used to calculate change in seal inner radius and final static,room temperature clearance.

The brush seal had a final clearance of 34.5 microns and the finger seal final clearance was 724 microns, which is 21 times greater than the brush seal.

---------------

Brush seal

• Calculated radius change

• Final clearance

131 µm

34.5 µm

Finger seal

889 µm

723.9 µm

NASA/CP—2005-213655/VOL1 127


Rotor Wear

• Rotor wear due to brush seal was very minimal.

• Rotor wear due to finger was measurable, but small.

Average rotor wear track measurements for finger seal

Rotor wear was measured using a profilometer at 8 circumferential locations around the rotor od to measure wear track depth and width.

The average wear track measurements for the finger seal are shown here and are quite small with a maximum track depth of about 6 microns.

Rotor wear due to the brush seal was minimal.

The Chrome Carbide coating performed well.

NASA/CP—2005-213655/VOL1 128


Conclusions

• Brush and finger seals have substantiality less leakage than labyrinth seals.

• The finger seal exhibited a more pronounced pressure closing effect than thebrush seal.

• Finger seal flow factor was 36% less than brush seal after 4 hour endurance test.

• Finger and brush seals have very similar power losses.

• Additional testing with cylindrical or labyrinth seal needed to establish pressurized tare data.

• Finger seal wear resulted in final radial clearance 21 times the brush seal.

• CrC rotor coating wear was minimal.

Self-Explanatory.

NASA/CP—2005-213655/VOL1 129


Lox Seal Test Rig During Test

This photograph shows the Lox Seal Test Rig during a test at the old Rocket Engine Test Facility.

As you may recall our cryogenic seals testing has been on hold due to the airport expansion.

Our cryogenic component laboratory has been re-located to Plum Brook Station in Sandusky, Ohio. The facility systems verification testing is beginning as we speak and should be completed by March 1, 2005. The facility has LN2, LH2, LOX, GN2, GH2, and GHe capabilities. We are working to get resources to re-establish the cryogenic seal test rigs at Plum Brook. If you have any need for cryogenic seals testing or testing of other cryogenic components, please let me know. We’d like to put this new facility to good use. You can contact me, Margaret Proctor, at 216-977-7526 or speak with me over lunch or at a break while you are here.

NASA/CP—2005-213655/VOL1 130

at Lewis FieldGlenn Research CenterUEET – Tip Clearance Control

Test Rig For Evaluating Active Turbine Blade Tip Clearance Control Concepts: An Update

Scott B. LattimeOhio Aerospace Institute, Cleveland, Ohio

Bruce M. Steinetz and Kevin J. MelcherNASA Glenn Research Center, Cleveland, Ohio

Jonathan A. DecastroQSS, Inc., Cleveland, Ohio

Malcolm G. Robbie and Arthur H. ErkerAnalex Corp., Cleveland, Ohio

2004 NASA Seal/Secondary Air System WorkshopNovember 9, 2004

Cleveland, Ohio

TEST RIG FOR EVALUATING ACTIVE TURBINE BLADE TIP CLEARANCE CONTROL CONCEPTS: AN UPDATE

Scott B. Lattime

Ohio Aerospace Institute Cleveland, Ohio

Bruce M. Steinetz and Kevin Melcher


Cleveland, Ohio

Jonathan A. Decastro QSS Group

Cleveland, Ohio

Malcolm G. Robbie and Arthur H. Erker Analex Corporation

Cleveland, Ohio

NASA/CP—2005-213655/VOL1 131


Objectives

• Design a fast mechanical ACC system for HPT tip seal clearance management– Improve upon slow thermal response of “case cooling” methods used today.– Provide continuous ACC throughout entire flight profile.

• Design a test rig to evaluate ACC system concepts.– Evaluate actuator concept response and accuracy under appropriate thermal and

pressure conditions in a non-rotating environment.– Evaluate clearance sensor response and accuracy in a non-rotating “hot“

environment.– Measure secondary seal leakage due to segmented shroud design, shroud

actuation, and case penetration.

• Test Rig Capabilities:– Chamber temperatures up to 1500 °F.– Seal carrier backside pressure up to 120 psi (simulate cooling air Δp).– Sized for actual seal hardware (20” diameter turbine).– Simulate realistic tip seal clearance changes due to mechanical and thermal

loading (electronically).– Positioning feedback sensing, rig construction, and actuation system designed to

achieve positioning accuracy 0.004-in.

NASA/CP—2005-213655/VOL1 132


Test Rig/ ACC System Specifications

• Temperature– The backside of the HPT shroud (blade

outer-air-seal) is generally cooled with compressor discharge air (P3 air: 1200 to 1300 ºF)

• Pressure– The cooling air is also used to purge the

leading and trailing edges of the shroud segments, providing a positive backflow margin from the hot rotor inlet flow.

– P3 is highest during maximum thrust events such as takeoff and re-accel. For large commercial engines this translates to a maximum cooling air pressure differential of up to 150-psid across the shroud.

• Actuation Range and Rate– Maximum tip clearance changes due to

axisymmetric and asymmetric loads on the rotor and stator components are on the order of 0.050-in.

– FAA requires that engines have the ability to reach 95% rated takeoff power from flight idle (or from 15% rated takeoff power) in 5.0 seconds.

→ 0.01-in/s

HPT shroud hanger and seal w/ cooling and purge flow

compressor discharge air

leading edge cooling holes

impingement cooling holes

trailing edge cooling holes

~0.7P3

~0.3P3

~0.8P3

P3

shroud hanger

shroud tip seal

turbine blade

1200-1300 °F

The design was concentrated on simulating the temperature and pressure conditions that exist on the backsides of the seal segments, without the need for a rotating turbine. This greatly simplified the rig design. We plan to assess the response of the ACC system to the effects of a turbine wheel (i.e., rapid clearance closures due to mechanical and thermal loads) by simulating closures electronically, as will be discussed in a later.

NASA/CP—2005-213655/VOL1 133


Design Criteria

• The substantial diameter of the segmented shroud structure (~20-in), under moderate pressures (~120-psi), gives rise to significant loads, and hence stresses, to which the actuation system and components must react.

• These stresses coupled with high temperatures (1200 to 1300 °F) can significantly reduce component cycle life due to creep.

• Managing these stresses with adequate materials and geometry to improve component cycle life was a driving factor in the rig component design.

• Larson-Miller parameter data for a variety of high temperature, super alloys was utilized to design components to achieve a desired minimum cycle life.

– Inconel 718 utilized for most of the hot section components.

– Components were designed for less than 0.5% creep strain, resulting in a 15-ksi limiting stress.

– 15-ksi stress level corresponds to over 100,000 hours life at 1300 ˚F and approximately 300 hours life at 1500 ˚F.

NASA/CP—2005-213655/VOL1 134


ACC Concept Evaluation Rigchamber

lower housing

radiant heater (lower half)

actuator mount

seal carrier assembly

air supply pipe (3 locations)

hydraulic servo actuator

Here we see the Gen 1 ACC System Concept and Test rig.

The test rig comprises six main components: the housing, the radiant heater, the pressurized chamber, the seal carrier assembly, the actuator rod assemblies, and the hydraulic actuators.

At the heart of the rig is a segmented shroud structure (seal carrier) that would structurally support the tip seal shroud segments in the engine. Radial movement of the seal carriers controls the effective position/diameter of the seal shroud segments, thereby controlling blade tip clearance.

The rig housing consists of two concentric cylinders, which form an annular cavity. An annular radiant heater made of upper and lower halves surrounds the segmented seal carrier structure to simulate the HPT tip seal backside temperature environment. A pressurized chamber encloses the carrier segments inside the annular heater through which heated pressurized air is supplied to simulate the P3 cooling/purge air pressure on the seal backsides. Heated air enters the chamber via three pipes that are fed from a manifold at the air heater exhaust through radial inlet ports as shown.

NASA/CP—2005-213655/VOL1 135


ACC Rig Cutaway

radiant heater

inlet air (Phigh)

exhaust air (Plow)

chamber

seal carrier

proximity probe

footactuator rod

main housing


movement


chamber air TC

flow deflector

actuator mount

The carrier segments are connected to independent hydraulic actuators through an actuator rod assembly. The foot of the actuator rod assembly positions the carrier segments in the radial direction, while allowing relative circumferential movement or dilation of the seal carrier segments through a pinned and slotted arrangement.

A series of radial tubes projecting outward from the chamber’s inner and outer side walls serve as supports, air supply and exhaust ports, probe fixtures, and the actuator rod guides. The chamber functions to support and align the carrier segments and actuator rods, as well as to house instrumentation and to seal the pressurized air from the radiant heater which is not designed to carry any pressure loading.

The inlet flow is baffled circumferentially around the outer chamber wall by a flow deflector to achieve uniform heating of the seal carrier assembly. The pressurized air is sealed along the sides of the seal carrier segments by contacting face seals that are energized via metal “E”-seals imbedded in the upper and lower chamber plates. The joints between adjoining carrier segments are sealed with thin metal flexures. Air that escapes over and between the carrier segments is vented out of the rig through a number of exhaust pipes that protrude radially along the inner chamber wall. The number and inner diameter of exhaust pipes were chosen to eliminate the possibility of backpressure at the exhaust ports.

High temperature proximity probes measure the radial displacement of the seal carriers at various circumferential locations. These measurements provide direct feedback control to the independent actuators and allow the desired radial position (clearance) to be set. The direct feedback control system allows for simulation of realistic transient tip clearance changes in lieu of a rotating turbine wheel. Superimposing a mission-clearance-profile over the actual clearance measurement input to the actuator controllers will allow researchers to assess the system’s response to the most dramatic transient events such as mechanical and thermal loading of the rotor during takeoff and re-accel.

The proximity probes are held at a constant standoff to the chamber inner wall via a spring-loaded piston arrangement. The spring-loaded mounting keeps the proximity sensor at a relatively constant position to the chamber inner sidewall during the initial heating of the rig. This arrangement also allows for easy removal of the probes without dismantling the housing.

The chamber air temperatures will be measured at three circumferential locations on the high-pressure side of the carriers to show how well the pressurized air is mixed by the chamber baffle. The chamber flange metal temperatures will be measured via two surface thermocouples attached to the inner and outer flange on the lower cover plate.

NASA/CP—2005-213655/VOL1 136


ACC Rig Test Stand

7’ 4’

Air Inlet

Air Exhaust

Air Heater

3’

Exhaust mixing air inlet

This slide shows the rig’s air supply and exhaust plumbing layout as well as the test stand dimensions.

The air heater system comprises two 36-kW, inline, flanged heaters, manufactured by Osram Sylvania. It is designed to supply up to 75-scfm of air at 120-psi and 1500 °F.

NASA/CP—2005-213655/VOL1 137


Carrier-Actuator Rod Assembly

• The slots are cut on a tangent to the radius on which the carrier pinholes are located.

• Prevents the carrier segment from cocking while it is displaced radially and circumferentially.

• The length of the slot allows a radial displacement of 0.2-in for each of the nine segments.

pivotslot

&10.25”

0.25”

2.0”

carrier pin(Inconel x750)

foot(Inconel 718)

seal carrier(Inconel 718)

flexure seal(Inconel x750)

actuator rod(Inconel 718)

Because the carriers are constrained by a pinned connection at one end and a slotted connection at the other, the segments must shift circumferentially as they are displaced in the radial direction. The slots in the feet are cut on a tangent to the radius on which the carrier pinholes are located. This keeps the carrier segments from cocking while it is displaced in both the radial and circumferential directions. The circumferential length of the carrier segments as well as the length of the slot in the actuator rod foot allows a radial displacement of 0.2-in for each of the nine segments. The slots for the flexure seals have adequate clearance to prevent the segments from becoming arch-bound as the segments are moved radially inward.

The pins are made of Inconel X750. This material was selected to help minimize galling against adjacent Inconel 718 components. The pins have flats machined on the diameter that contacts the slots, providing a bearing surface and reducing the contact stress developed between the pin and foot

NASA/CP—2005-213655/VOL1 138


Cooled Actuator Rod Design

exit cooling holes air inlet

air outlet

actuator mount

inlet cooling holes

ring seals (Stellite 25 /Inconel 625)

elastometric O-rings

support tube

actuator rod

The cooling scheme allows the actuator rod and support tube to function as a tube-in-tube heat exchanger using a small flow rate of ambient air to cool the assembly.

The cooling holes were made from three sets of six, 0.03-in diameter holes drilled around the circumference of the hollow rod. Ambient air, supplied at the rod end through features in the actuator mount, travels axially through the center of the rod, passes radially through the cooling holes, and exits between the support tube and outer diameter of the rod.

NASA/CP—2005-213655/VOL1 139


Rod Cooling Scheme Validation

0

200

400

600

800

1000

1200

1400

1600

8:24 9:36 10:48 12:00 13:12 14:24 15:36 16:48

time

tem

pera

ture

(8F

)0.0

0.5

1.0

1.5

2.0

2.5

3.0

3.5

4.0

4.5

flow

(sc

fm)furnace

masstube endrod endflow (scfm)

test hardware in furnace

air inletrod

tubemass (foot)

Air flow of 4.0 scfm kept both rod and tube ends < 250 °F.

A mockup of the cooled actuator rod design was built to validate the design. A solid steel block (simulating the actuator foot) was bolted to one end of a stainless steel tube (simulating the rod). Another larger tube was welded to the block (simulating the support tube) in a concentric arrangement. An air supply line was attached to the end of the inner tube from which the assembly was supported and inserted into a box furnace. The insulation thickness of the furnace closely approximated that of the radiant heater designed for the rig. A plastic supply line was used minimize heat loss through the supply tube. Thermocouples were attached to measure the temperatures of the mass, the end of the inner rod, and the end of the outer tube. The furnace was heated to 1500 °F and after the mass temperature stabilized at 1500 °F, ambient air at approximately 70 °F was supplied to the assembly. Temperatures of the furnace, mass, tube end, and rod end are shown as a function of time on the left vertical axis. The cooling air volumetric flow is shown on the right vertical axis. The chart shows that for minimal air flow (3.0 to 4.0-scfm) both the tube and rod end temperatures were kept below 250 °F. Thus the cooling scheme design was successfully validated and implemented into the rig design to allow the use of conventionalactuators.

NASA/CP—2005-213655/VOL1 140


Chamber Seal Locations & Leakage

1.611500120C-seal (outer flange)

5.401500120E-seals (behind face)

2.101500120face seals

0.651500120piston rings

33.1Total

23.341500120flexure seals

Flow(scfm)

Temp (°F)

ΔP (psi)

location

E-seal (Waspaloy)

face seal (Stellite 6B)

flexure seal (Inconel X750)

sacrificial stud (Inconel 718)

Phigh ~ 120-psig T ~ 1500 °F

bearing pads

C-seal (Waspaloy)

Plow ~ 0-psig T ~ 1500 °F

ring seals (Stellite 25 /Inconel 625)

The nature of this ACC concept with its segmented shroud design and case penetration requires multiple high temperature seals. The test rig will allow the development and evaluation of these seals that will eventually be required in an engine. Obviously the leakage created by the use of these secondary seals must be minimized to gain the benefits of the ACC system. For the test rig, the secondary seal leakage drove the design of the air heater.

Flexure seals are used to prevent the radial flow of pressurized air between the carrier segment joints. The 2.0-in wide by 0.9-in long flexures are made of 0.02-in Inconel X750 sheet stock. This material was chosen for its galling resistance to the carrier material. The carrier slits that contain the flexures are designed with a 0.01-in clearance.

The chamber contains four “C-seals”, two on the upper and lower outer diameter flanges and two on the upper and lower inner diameter flanges of the cover plates. The seals are made of Waspalloy and have a cross sectional thickness of 0.015-in. The seals were designed by Perkin-Elmer to seal against a 120-psi pressure at 1500 °F and they require a 150-lbf/in seating load per seal at assembly.

The upper and lower cover plates also house a metal face seal assembly. These seals act to block the pressurized air from flowing between the cover plates and carrier segments. The face seal, made of Stellite 6B, is a pressure balanced design and utilizes an “E-seal” as a preload and secondary seal device. Stellite 6B was selected for the face seal material due to its high temperature properties and anti-galling performance against Inconel 718. The E-seals, also designed by Perkin-Elmer Fluid Sciences and made of Waspalloy, provide a closing force to the face seal on the carrier segments and prevent air from leaking between the face seal and cover plate. Each E-seal provides about 10-lbf/in preload to its corresponding face seal. The face seal was designed with a generous cross section, due to its large diameter, to provide stiffness for operation as well as manufacturing.

The actuator rod contains two pairs of expanding concentric ring seal sets on its bearing surface. Each pair is made of an outer Stellite 25 ring and an inner Inconel 625 ring. The seals were designed by Precision Rings, Inc. (Indianapolis, IN) for a 120-psi at 1500 °F.

NASA/CP—2005-213655/VOL1 141


ACC Controller Framework

ActuatorPositionSensor

++

ACC Test Rig

Real-time Controller& Simulation

Desired Clearance

SimulatedTransients

+-

ShroudActuator

Error Shroud PositionActuator

Command

Actuator Position

SimulatedClearance

AdvancedMultivariableControl Laws

ClearanceProbe

Engine Model

This slide shows a diagram of the control system strategy that will employed to operate and evaluate this first generation ACC system as well as future systems. Each of the nine independent hydraulic actuators will have its own feedback control allowing the positioning of each cylinder or actuator. An outer loop will be monitoring the position of the carrier segments. The outer loop will determine the minimum clearance off of which the desired clearance will be measured. The control system will be used to evaluate the accuracy and response of the ACC system to both steady state and transient clearance profiles.

Our next speaker, Mr. Kevin Melcher of the Controls and Instrumentation Branch at NASA GRC, will provide a more in depth discussion on his development of this control system and his work on defining control system requirements and architecture for advanced ACC systems.

NASA/CP—2005-213655/VOL1 142


Test Rig Fabrication Status

Housing with lower heater half & TC’s

ACC Rig with shroud assembly

Shroud Assembly with Flow Deflector

capacitance sensor

ACC Rig with Chamber Installed

Displacement Sensor Assembly

Housing with upper & lower heater halves, TC’s

Here we can see some of the main components of the test rig as they are currently being fabricated. The components are about 75% complete. We expect assembly to occur at the end of the month, with rig check out occurring towards the end of December.

NASA/CP—2005-213655/VOL1 143


Rig Assembly w/ Mechanical Checkouts & Gage Plate

Gage Plate

Mechanical Checkout

NASA/CP—2005-213655/VOL1 144


Carrier Segments Cycled Through 0.2”

Gage Plate Pin Captured at Neutral Diameter

Segments at + 0.1”( &max )

Segments at - 0.1”( &min )

NASA/CP—2005-213655/VOL1 145


Chamber Hydro Test

Deflections measured over face seal groove

0

2

4

6

8

10

12

14

0 100 200 300 400

psig

tho

san

dth

s o

f an

inch

Cracked boss weld

NASA/CP—2005-213655/VOL1 146


Schedule

Q1 FY05

Q1 FY05

Q2 FY05

Q2 FY05

Q1-2 FY06

Q1-2 FY06

Complete Assembly/ Facilities

Complete Rig Checkout

Evaluate Gen 1 ACC Concept (Ambient)

Evaluate Gen 1 ACC Concept (~1300 °F)

Evaluate AADC SMART Track Concept (Ambient)

Evaluate AADC SMART Track Concept (~1300 °F)

NASA/CP—2005-213655/VOL1 147


Plans

• The ACC Test Rig will be used to evaluate the performance of GRC’s 1st Gen ACC system and others (AADC –SMART Track). Performance will be evaluated under a series of HPT simulated temperature and pressure conditions. – Actuator stroke, rate, accuracy, repeatability– System concentricity, synchronicity– Component wear– Secondary seal leakage– Clearance sensor response and accuracy

• The results of this testing will be used to further develop/refine the current actuator designs as well as other actuator concepts. – SMA’s, piezoelectric, magnetostrictive, other

• Optimization of ACC system components for future flight hardwaredevelopment.– increased cycle life – reduced size and weight

NASA/CP—2005-213655/VOL1 148

NASA Seals WorkshopNovember, 2004

NASA Seals WorkshopNASA Seals WorkshopNovember 9November 9--10, 200410, 2004

Norman Turnquist – GE Global Research

Farshad Ghasripoor – GE Global Research

Mark Kowalczyk – GE Energy

Bart Couture – GE Energy

Wear Prediction of Strip Seals through Wear Prediction of Strip Seals through ConductanceConductance

WEAR PREDICTION OF STRIP SEALS THROUGH CONDUCTANCE

Norman Turnquist and Farshad Ghasripoor GE Global Research

Niskayuna, New York

Mark Kowalczyk and Bart Couture GE Energy

Schenectady, New York

NASA/CP—2005-213655/VOL1 149

Advanced Sealing Technologyat GE Global Research

TeamShorya AwtarNitin BhateBruce BrielBruce BrissonRay ChuppBiao FangFarshad GhasripoorChuck GoldenMohsen SalehiOmprakash SamudralaNorman TurnquistKripa VaranasiChris Wolfe

Advanced Seal team at GE Global Research

-Developing advanced seal technologies for GE Aircraft Engines, Land-Based Gas Turbines, Steam Turbines, Industrial Compressors, etc.

-Developing seals for both new products and retrofits into existing equipment

NASA/CP—2005-213655/VOL1 150


Advanced Sealing Synergy

• Rotor dynamics• Frictional heating• Rub tolerant seal

• Short Cycle• Low Cost

• Multi-Stage

• Seal Stability• High swirl ratio & High speed• High Temps & Creep• Seal life & Reliability• System Integration

• Longer life: 48,000 hrs• Large Interference• Secondary flow

system optimization• Field Performance

Monitoring

Gas Turbine

Steam Turbine

Aircraft Engine

GRCSeal fundamentals

GE AE/PS applicationsSeal Design/Development/Testing

Generator• Oil & H2 Sealing

• Non-metallic Seals

Compressor• Reverse Rotation

• Particles • Chemicals• 2500 psi

Seal design toolsSeal wear/life

StabilityReliability

Field Performance & Validation

Different applications have different design considerations for advanced seals.

Seals are developed for each application, analytical design tools are developed, and the design tools can be applied across the spectrum of applications.

GE Global Research is a “hub” for Advanced Seals development across GE.

NASA/CP—2005-213655/VOL1 151


High Pressure RigStatic Seal Rig

Large Diameter Rig Abradable Seal Rig

ADVANCED SEALS - Test Facilities at GE-GRC

Several experimental rigs are used to quantify performance and characteristics of advanced seals. A static rig is employed for testing both static and dynamic seals. It is a high pressure, high temperature rig that gives comparative leakage performance data for various seals types, i.e., cloth, labyrinth, brush, honeycomb, C-, E-, etc. A smaller rotary rig (High Pressure Rig) is used for testing in air or steam of dynamic seals. This rig is used to test subscale seals at full turbine conditions. A larger rotary rig is used for testing full-scale dynamic seals at subscale conditions. This rig is the one that has been used to test aspirating face seals as well as brush seals. An abradable rub rig is a versatile rig for testing candidate abradable shroud materials rubbing against tipped and untipped blades. It can simulate turbine blade-tip rotation and incursion rates, and has heating capability to operate at turbine environment temperatures. Wear characteristics are determined from measured level of shroud and blade-tip wear.

NASA/CP—2005-213655/VOL1 152


Abradable Abradable -- Strip Seal Strip Seal ConfigurationConfiguration

Help reduce clearances at tips:

• 2% Section Efficiency

• Easily refurbishable

Detail of strip seal configuration

GE HEAT Steam Turbine has 28 stages.

Caulked-in strip seals used in drum rotor turbine construction; strips caulked directly into rotor.

Strips used in conjunction with abradable coating on shrouded nozzle tips.

Strips can be replaced if damaged.

Abradable coating permits tight assembly clearances without risk of hard metal-to-metal rubs.

NASA/CP—2005-213655/VOL1 153


Mushroomed vs. UnMushroomed vs. Un--Mushroomed strip after rubsMushroomed strip after rubsagainst an abradable coatingagainst an abradable coating

Broken strip after a Broken strip after a hard rubhard rub

Able to simulate rub events using a lab based rig

Rig Capability:• Surface Speeds up to 335 m/s (1100 ft/s)• Temperatures up to 1000 °C (1832) °F• Radial, Axial and combined incursions between

stator & rotor• Incursion rates from 0.5 μm/s – 38 mm/s

(0.00002 – 1.5 in/s)• Continuous rails/strips or individual blades

Subscale rub tests demonstrate effectiveness of abradable coating in preventing strip damage during rub events test rig also used to investigate rub behavior of various strip materials.

NASA/CP—2005-213655/VOL1 154


The combined effect of high temperature increasing electron energy and electron scatter can lead to very different thermal behaviors for different materials

Thermal Conductivity Thermal Conductivity & Conductance& ConductanceTHERMAL CONDUCTANCETHERMAL CONDUCTANCE - Also known as Heat Transfer Coefficient is a measure of the rate at which heat energy flows through a surface. It is a measure of the amount of energy flowing through a unit area, in unit time, when there is a unit temperature difference between the two sides of the surface.

Strip tip-base temperature disparity during rub

Definition of Aspect Ratio (α): (Thk/H)

.000"R

.000" Seal

.000 +.000 -.000

.000" +/-.000

.000"

.000"

Square Caulk Wire

t2

P

ra

rb

>900 °C

~200 °C

Conductance accounts for strip geometry as well as conductivity of the strip material. Ability of strips to draw away heat from rub event is key to minimizing strip wear.

Aspect ratio is an important variable that is purely a function of the strip geometry.

NASA/CP—2005-213655/VOL1 155


Conductance as a function of aspect ratioConductance as a function of aspect ratio

Increasing strip thickness and/or aspect ratio improves conductance

Conductance increases linearly with increasing aspect ratio.

NASA/CP—2005-213655/VOL1 156


Wear of strip as a function of strip aspect ratioWear of strip as a function of strip aspect ratio

Un-Mushroomed tip

Mushroomed tip

• Increasing strip thickness improves wear behavior• Strip wear is minimum for aspect ratios >0.07

Increasing strip thickness increases Aspect Ratio, and therefore Conductance.

For the cases considered, Aspect Ratio = 0.07 is the cutoff between excessive and acceptable strip wear.

NASA/CP—2005-213655/VOL1 157


Wear of strip as a function ConductanceWear of strip as a function Conductance

• Transition in wear behavior occurs over a narrow range of conductance• Beyond a certain conductance (i.e. aspect ratio) conductance model for

rub is ineffective

The transition from Severe Wear to Mild Wear is abrupt, and occurs when the heat transferred away by Conductance is inadequate to prevent melting at the tip.

NASA/CP—2005-213655/VOL1 158


Thermal Conductivity vs. Tensile StrengthThermal Conductivity vs. Tensile Strength

• Strength of strip has secondary effect at rub surface speeds in excess of 50m/s• Conductivity has the most effect on rub behavior of strip

Tensile strength of the strip material has only a secondary effect on wear behavior; the wear is primarily driven by conductance.

The implication is that high strength superalloys are not necessarily good strip materials from a wear standpoint.

NASA/CP—2005-213655/VOL1 159


Predictive wear modelPredictive wear model

Mathematical correlation is established based on test data

A predictive model for strip wear has been developed and validated with subscale rig tests.

NASA/CP—2005-213655/VOL1 160


<418.0°F

1268.1°F

AR01

Test Test –– High thermal Conductivity stripHigh thermal Conductivity strip

Tested strip aspect ratio

Baseline temperature of ~ 350 °C with peaks approaching ~700 °C.

Strip temperatures measured using thermal imaging show peak temperatures of ~700C for a high conductivity strip

NASA/CP—2005-213655/VOL1 161


Test Test –– Low thermal Conductivity stripLow thermal Conductivity strip

150.7°F

1874.6°F

1000AR01

Strip temperature during rub validates conductance model

Tested strip aspect ratio

Baseline temperature of ~ 550 °C with peaks approaching ~1000 °C.

Strip temperatures measured using thermal imaging show peak temperatures of ~1000C for a low conductivity strip.

NASA/CP—2005-213655/VOL1 162


ConclusionsConclusions

Test data indicate strong influence of conductance in rub behavior of strip seals at speeds in excess of 50 m/s.

Material strength appears to have little effect on rub for strip seals with aspect ratios <0.07

A predictive model is established as a tool to help select the strip material and in design of the strip

Best strip material selection and design would extend the strip seal’s ability to severe rub events beyond the abradable thickness

Conclusions

NASA/CP—2005-213655/VOL1 163

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CENTURION TM

MECHANICAL SEALS

CenturionCenturionTMTM Mechanical SealsMechanical Seals

Parametrical Study of Hydrodynamic Parametrical Study of Hydrodynamic Seal Using a 2D Design Code and Seal Using a 2D Design Code and

Comparing with 3D CFD ModelComparing with 3D CFD Model

Contact Info:

Xiaoqing ZhengCenturion Mechanical Seals

15 Pioneer AvenueWarwick RI 02888

(401) [email protected]

PARAMETRICAL STUDY OF HYDRODYNAMIC SEAL USING A 2D DESIGN CODE AND COMPARING WITH A 3D CFD MODEL

Xiaoqing Zheng

Perkin Elmer Centurion Mechanical Seals Warwick, Rhode Island

Hydrodynamic film-riding gas seal, having been successfully used in industrial pump and compressor applications for decades, is now finding way into aero-engines. Aerospace applications require the seal work robustly under various speeds and constantly changing pressures. Maintaining seal face flatness is also more challenging in aerospace than in industrial applications because of the seal size and the rapid thermal transition in aero-engines. Another difficulty in aerospace application comes from the altitude condition, where air is thin, not enough opening force may be generated to separate the seal faces from touching. Therefore, the hydrodynamic groove design is more critical in aerospace application. In order to maximize the seal performances, the groove shape and depth are optimized for the worst application condition, such as at the altitude, with using a 2D model coupled with a non-linear optimization procedure.

Seals designed in this way have been found working well both in rig test and in on-flight evaluation. However, two recently designed seals based on full optimization were found not working during rig test. These two seals features very fine grooves that is thinner than the other seals that worked well. The 2D code from time to time predicted that seals work better if more grooves were engraved. The under-performed two seals indicted there might be something missing in the 2D model. It was speculated that the entrance losses and 3D effects for narrow grooves become significant, resulting in reduction of hydrodynamic opening force. In order to quantify the influences of 3D effect, a three-dimensional model is therefore developed to calculate the fluid flow between the seal faces, and then the results are compared with the 2D solution. The model solves a circumferential section of seal face that consists one land and one groove region. A small region outside the sealing face is also included to simulate the flow entrance and exit. It was found that the 3D CFD model generally agrees with the 2D model in determining the optimum groove dimensions. The 3D code gives similar trend predictions of seal opening force under the influences of groove depth, number of grooves, land to groove ratio and groove angle, although there exist differences in the absolute value levels. The 3D CFD results forced a re-examination of hardware and further re-test the seal. It is concluded that the problems were due to manufacture tolerance with such fine grooves. The design code is correct, but manufacture capability should be incorporated. CFD results and test data are reported in this paper.

NASA/CP—2005-213655/VOL1 165

CENTURION TM

MECHANICAL SEALS

IntroductionIntroduction

Hydrodynamic Face Seal• Riding on a thin film• No contacting, no wear• Extremely small air leakage, no oil leakage

Challenge in Aerospace Application• High-altitude: Thin air• Transient• Large diameter

Hydrodynamic film-riding gas seal, having been successfully used in industrial pump and compressor applications for decades, is now finding way into aero-engines. Aerospace applications require the seal work robustly under various speeds and constantly changing pressures. Maintaining seal face flatness is also more challenging in aerospace than in industrial applications because of the seal size and the rapid thermal transition in aero-engines. Another difficulty in aerospace application comes from the altitude condition, where air is thin, not enough opening force may be generated to separate the seal faces from touching. Therefore, the hydrodynamic groove design is more critical in aerospace application. In order to maximize the seal performances, the groove shape and depth are optimized for the worst application condition, such as at the altitude, with using a 2D model coupled with a non-linear optimization procedure.

NASA/CP—2005-213655/VOL1 166

CENTURION TM

MECHANICAL SEALS

1. Seal Equation:

2. Code uses CFD to solve groove equations

3. Code is coupled with ADINA to analyze face deflections

4. Optional Dam Section Hydrostatic Equation: (for high pressure drop)

PKI “State Of The Art” Design Code Solves The Seal Equations

0}6{ 3 =−∇•∇ θωμ rphpphrr

hM

fMM

hdr

dh

rM

MM

dr

dM

)1(

)1(2)

1(

1

)1(22

22

14

2

222

12

−+++

−+=

−− γγ γ

Film Riding Hydrodynamic Face SealFilm Riding Hydrodynamic Face Seal

Automatic 6 design Automatic 6 design parameters Nonparameters Non--linear linear optimizationoptimization

The 2D seal code has been validated against rig test data including high-altitude conditions. It is then incorporated into the powerful ADINA program for 3D dynamic analysis. A non-linear procedure has been developed on top of the program. Each design is optimized based on the worst operating condition. Therefore each design is a result of full DOE analysis with direct simulations. As shown here, each design is distinctive in the shape, depth, and number of grooves.

NASA/CP—2005-213655/VOL1 167

CENTURION TM

MECHANICAL SEALS

Design Tool ReliabilityDesign Tool Reliability？？

• When seal failed• When seal contacts / wears• When certain kind of seal under-performed

Is the design tool reliable?If not sure, what is missing?

❧ 2D simplification in fluid simulation?❧ Entrance loss (inadequate computational domain)?

Seals designed in this way have been found working well both in rig test and in on-flight evaluation. However, two recently designed seals based on full optimization were found not working during rig test. These two seals features very fine grooves that is thinner than the other seals that worked well. The 2D code from time to time predicted that seals work better if more grooves were engraved. The under-performed two seals indicted there might be something missing in the 2D model. It was speculated that the entrance losses and 3d effects for narrow grooves become significant, resulting in reduction of hydrodynamic opening force.

NASA/CP—2005-213655/VOL1 168

CENTURION TM

MECHANICAL SEALS

Design / Analytical ToolsDesign / Analytical Tools

• ADINA - Integrated Structural, Thermal, Dynamic and CFD Analysis

• Multiple Proprietary Design Codes (ADINA FEA plug-ins)

ADINA, a multi-physics engineering program specialized in Fluid-Structure Interaction simulation is used in our seal design and optimization. Uniquely, we have plugged our seal code into the ADINA system, therefore 3D CFD, Structural, Thermal dynamic analysis can be done efficiently for hydrodynamic seal design.

NASA/CP—2005-213655/VOL1 169

CENTURION TM

MECHANICAL SEALS

• 3D CFD hydrodynamic analysis

3D CFD Model3D CFD Model

3D Gas Film Surface Pressure

Grooves

Gas film

Gas film Carbon OD

Carbon face

Mating ring surface

Groove Side

Groove ID

Groove OD Carbon ID

Periodical boundary

Block Outside of the Sealing region

The fluid flow domain includes the groove that extends beyond the carbon ID, the land area, dam area and areas outside of the sealing faces to include entrance losses in the model. One slice consisting of one groove and one land area is actually solved.

NASA/CP—2005-213655/VOL1 170

CENTURION TM

MECHANICAL SEALS

Governing EquationsGoverning Equations

Governing Equations

• The equations solved are 3D Navier-Stokes Equations. No simplification is made.

Boundary conditions

1. The mating ring surface and grooves are rotating at a constant speed and the seal ring is stationary. The seal OD and ID are exposed to ambient pressure and room temperature.

2. Stationary and rotating walls are treated as non-slipping

3. Periodical boundary condition for section interfaces

NASA/CP—2005-213655/VOL1 171

CENTURION TM

MECHANICAL SEALS

Model SealModel Seal

Parameter Value Notes

Groove profile radius 1.47

Film thickness 0.000100

Number of grooves 20-70 While land to groove ratio is kept 0.75

Land to groove ratio 0.25-2 Number of groove is fixed

Groove depth 100-750 groove number and angle are fixed

Groove angle 6-30 Groove depth, number of grooves and land to groove ratio are fixed

Shaft speed 10,000 and 20,000 rpm

Pressure Ambient for ID and OD

NASA/CP—2005-213655/VOL1 172

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MECHANICAL SEALS

The Influences of Number of GroovesThe Influences of Number of Grooves

Influence of Number of Groovesat 20,000 rpm (260 ft/sec)

0

2

4

6

8

10

12

14

25 35 45 55 65 75

Number of Grooves

Net

Op

enin

g F

orc

e

AeroSeal 3D CFD

The influences of number of grooves for this seal are studied at 10,000 rpm and 20,000 rpm. The groove depth is set to 0.000350 inches more than the design value, the land to groove ratio is 0.75, and the groove angle is set to the design value. Both 2D and 3D models predict that more grooves, more opening force. Despite the difference of absolute opening force between the 3D and 2D results, the trend is parallel for both models.

NASA/CP—2005-213655/VOL1 173

CENTURION TM

MECHANICAL SEALS

The Influences of Groove DepthThe Influences of Groove Depth

Influence of Groove Depth

02468

101214161820

0 200 400 600 800

Groove Depth (micro-inch)

Hyd

rod

ynam

ic

Fo

rce(

lbf)

3D CFD (10000 rpm) AeroSeal (10000 rpm)


The groove number, the land to groove ratio and the groove angle are set to the design values in this study. Both 2D and 3D models show strong dependency of hydrodynamic force on the groove depth. It is also interesting to note the difference between 2D and 3D model becomes smaller at the groove depth gets shallower or getting very deep.

In this case, the optimal groove depth under seal level pressure condition is close to the design value, predicted by both 2D and 3D models

NASA/CP—2005-213655/VOL1 174

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MECHANICAL SEALS

The Influences of Groove AngleThe Influences of Groove Angle

Groove Angle Effects

0

2

4

6

8

10

12

14

16

18

20

5 10 15 20

Groove Angle (degree)

Dyn

amic

Op

enin

g F

orc

e



The groove depth and the land to groove ratio are fixed with design values. The groove radius of curvature is 1.470 inches representing a spiral groove at the initial angle of groove.

Both 2D and 3D models show considerable dependency of hydrodynamic force on the groove angle. In this study, the groove profile curvature is fixed; only angle is changed to isolate the influence. For real spiral grooves, the curvature is a function of the groove angle.

In this case, the optimal groove angle predicted by the 2D model is only one degree higher than that by the 3D model.

NASA/CP—2005-213655/VOL1 175

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MECHANICAL SEALS

The Influences of Land to Groove RatioThe Influences of Land to Groove Ratio

Land to Groove Ratio Effects

0

2

4

6

8

10

12

14

16

18

20

0 0.5 1 1.5 2 2.5

Land to Groove Ratio

Dyn

amic

Op

enin

g F

orc

e



NASA/CP—2005-213655/VOL1 176

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MECHANICAL SEALS

Revisit Tested SealRevisit Tested Seal

❖ Both 2D and 3D models suggest more grooves work better within the range that we are looking

Manufacture Discrepancy

➘ The designed land to groove ratio is 0.75. ➘ The made is only 0.5. The grooves were made too

wide. ➘ If the land to groove ratio is set to 0.5, rerun the

2D seal analysis confirms rig observation.

NASA/CP—2005-213655/VOL1 177

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MECHANICAL SEALS

ReRe--TestTest

0

50

100

150

200

250

300

350

0 20 40 60 80 100 120

Time

Tem

per

atu

re(F

)

0

5

10

15

20

25

30

35

Sp

eed

(Krp

m),

Pre

ssu

re(p

si)

Air Cavity Seal Temp. #1 Seal Temp #2 Oil inOil out Speed (krpm) Air Pressure

Temperature increases as pressure

Seal touches down

Seal lifts off

NASA/CP—2005-213655/VOL1 178

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MECHANICAL SEALS

ConclusionConclusion

1. The 3D CFD model generally agrees with the 2D model.

2. The physical model of the 3D code is more accurate than that of the 2D code.

3. However, the 2D model is solved numerically more accurately than the 3D model is.

4. In most cases, At least theoretically, seals with more grooves work better.

5. 3D CFD simulation also supports the groove design from the 2D model.

NASA/CP—2005-213655/VOL1 179

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MECHANICAL SEALS

Production Seal ExampleProduction Seal Example

Lift Off Grooves Clean

920 Hours on Field Evaluation Seals (Jan 2004)920 Hours on Field Evaluation Seals (Jan 2004)

This is a picture we obtained from Honeywell earlier this year. It shows the hydrodynamic grooves still clean and there is no oil leakage. The seal has been qualified for production and has accumulated more than 51,000 hours in field with no failure.

NASA/CP—2005-213655/VOL1 180

ATIGAT

NON-CONTACTING FINGER SEAL DEVELOPMENTS AND DESIGN CONSIDERATIONS:

Thermofluid & Dynamics Characterization, Experimental

M.J. Braun, ProfessorH. M. Pierson, D. Deng, Graduate StudentsDepartment of Mechanical Engineering,University of Akron, Akron, Ohio

M. Proctor, NASA GRC, Cleveland, OhioTechnical Supervisor

NASA Seal Workshop, November 9-10, 2004Cleveland, Ohio

NON-CONTACTING FINGER SEAL DEVELOPMENTS AND DESIGN CONSIDERATIONS: THERMOFLUID AND DYNAMICS CHARACTERIZATION, EXPERIMENTAL

M. Jack Braun, Hazel M. Pierson, and Dingeng Deng

University of Akron Akron, Ohio

Margaret P. Proctor


Cleveland, Ohio

NASA/CP—2005-213655/VOL1 181

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NASA/CP—2005-213655/VOL1 186

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NASA/CP—2005-213655/VOL1 192

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NASA/CP—2005-213655/VOL1 193

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NASA/CP—2005-213655/VOL1 194

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-DO

F

⎭⎬⎫

⎩⎨⎧=

⎭⎬⎫

⎩⎨⎧ ⎥ ⎦⎤⎢ ⎣⎡

+−

−+

⎭⎬⎫

⎩⎨⎧ ⎥ ⎦⎤⎢ ⎣⎡

+−

−+

⎭⎬⎫

⎩⎨⎧ ⎥ ⎦⎤⎢ ⎣⎡

00

01

21

21

21F

xx

kk

k

kk

xx

cc

c

cc

xx

m

m

sf

f

ff

cf

f

ff

FS

R

&

&

&&

&&

()

yk

yc

xk

kx

cx

m

FE

quF

Equ

SEqu

FE

quF

Equ

Equ

+=

=+

++

&

&&&

Fin

ger

seal

mas

s fo

llow

ing

a di

spla

cem

ent

inpu

t

Inte

ract

ion

of fi

nger

and

rot

or m

ass

thro

ugh

the

flui

d la

yer

and

subj

ect

to a

forc

ing

func

tion

inpu

t +

Cou

lom

b fr

icti

on

NASA/CP—2005-213655/VOL1 195

AT

IGA

TA

TIG

AT

ti e

FF

ω0

= Now

con

side

ring

the

inpu

t for

ce to

be

harm

onic

,, w

e ca

n as

sum

e th

e re

spon

se o

f the

mas

ses

to b

e ch

arac

teri

zed

by

() ()t

i

ti e

tx

et

xωω

22

11

Χ=

Χ=

Thu

st

iti e

x

ei

xω

ω

ωω

Χ−

=

Χ=

22,1

2,1

&&

&

Subs

titu

ting

into

the

mat

rix

equa

tion

abo

ve

⎭⎬⎫

⎩⎨⎧=

⎭⎬⎫

⎩⎨⎧ ΧΧ⎥ ⎦⎤

⎢ ⎣⎡

ΖΖ

ΖΖ

00

21

2221

1211

iwt

eF

whe

re

()

()

sf

cf

FS

ff

ff

R

kk

cc

im

kc

i

kc

im

++

++

−=

Ζ

−−

=Ζ

=Ζ

++

−=

Ζ

ωω

ωω

ω

222

2112

211

The

sol

utio

n is

then

()

[]

Fi

vv

1−Ζ

=Χ

ω2-D

OF

co

nt’

d

NO

TE

: The

mot

ion

of th

e fin

ger

whe

n th

e ro

tor/

finge

r in

tera

ct is

rep

rese

nted

by

thes

e eq

uatio

ns a

nd s

olut

ion

NASA/CP—2005-213655/VOL1 196

AT

IGA

TA

TIG

AT

Fre

e vi

bra

tio

n o

f th

e fi

ng

er (

2 H

P-

2 L

P) w

ith

an

init

ial p

osi

tive

rad

ial d

efle

ctio

n a

nd

va

ryin

g a

mo

un

ts o

f C

ou

lom

b f

rict

ion

.

x+FS

MA

SSFIN

GE

RS

TIF

FN

ESS

BA

CK

PLA

TE

fF

kxx

m±

=+&&

The

+ a

nd –

sign

s ar

e de

pend

ent o

f the

dire

ctio

n of

mot

ion

of th

e fin

ger

and

thus

orie

ntat

ion

of th

e C

oulo

mb

fric

tion

for

for

nt

ωπ

≤≤

0 ()kF

tkF

xt

xf

nf

+⎥ ⎦⎤

⎢ ⎣⎡−

=ω

cos

0

nn

tω

πω

π2

≤≤

()kF

tkF

xt

xf

nf

−⎥ ⎦⎤

⎢ ⎣⎡−

=ω

cos

30

nn

tω

πω

π3

2≤

≤

()kF

tkF

xt

xf

nf

+⎥ ⎦⎤

⎢ ⎣⎡−

=ω

cos

50

for

NASA/CP—2005-213655/VOL1 197

AT

IGA

TA

TIG

AT

-1.5

000E

-03

-1.0

000E

-03

-5.0

000E

-04

0.00

00E+

00

5.00

00E-

04

1.00

00E-

03

1.50

00E-

03 0.00

000.

0010

0.00

200.

0030

0.00

400.

0050

0.00

600.

0070

0.00

tim

e (

s)

displacement (inches)

0 ps

i1

psi

2 ps

i3

psi

4 ps

i5

psi

6 ps

i7

psi

8 ps

i9

psi

10 p

si

Pre

ssur

e D

rop

thro

u gh

equa

lizat

ion

hole

s (d

elta

p)

Fre

e F

ing

er M

oti

on

Wit

h V

ario

us

Am

ou

nts

Of

Co

ulo

mb

Fri

ctio

n. F

ing

er S

tiff

nes

s is

Co

nst

ant

The

se le

vels

do

mot

hav

e en

ough

ene

rgy

left

to

mov

e ba

ck t

o eq

uilib

rium

line

The

se le

vels

do

mot

hav

e en

ough

ene

rgy

left

to

mov

e ba

ck t

o eq

uilib

rium

line

Cou

lom

b fr

ictio

n W

INS

Init

ial D

isp

lace

men

t o

f th

e fi

ng

er

NASA/CP—2005-213655/VOL1 198

AT

IGA

TA

TIG

AT

Fre

e F

ing

er M

oti

on

Wit

h C

on

stan

t C

ou

lom

b

Fri

ctio

n. F

ing

er S

tiff

nes

s is

Var

ied

.

∆p

=5 p

si

∆p

=10

psi

-0.0

01

-0.0

008

-0.0

006

-0.0

004

-0.0

0020

0.00

02

0.00

04

0.00

06

0.00

08

0.00

1

00.

001

0.00

20.

003

0.00

40.

005

0.00

6

Tim

e (

s)

Displacement from Equilibrium (in)

0.5k

(11

.72)

0.75

k (1

7.58

)

curr

ent

k (2

3.44

)1.

5k (

35.1

6)

2.0k

(46

.88)

2.5k

(58

.60)

3.0k

(70

.32)

Equ

ival

ent

Stic

k S

tiffn

ess

(lbf/

in)

-0.0

01

-0.0

008

-0.0

006

-0.0

004

-0.0

0020

0.00

02

0.00

04

0.00

06

0.00

08

0.00

1

0.00

12

0.00

14

0.00

16

00.

001

0.00

20.

003

0.00

40.

005

0.00

6

Tim

e (

s)Displacement from Equilibrium (in)

0.5k

(11

.72)

0.75

k (1

7.58

)

curr

ent

k (2

3.44

)1.

5k (

35.1

6)2.

0k (

46.8

8)2.

5k (

58.6

0)

3.0k

(70

.32)

Equ

ival

ent

Stic

k S

tiffn

ess

(lbf/

in)

As

the

stic

k be

com

es s

tiffe

r, th

e C

oulo

mb

fric

tion

has

less

effe

ct o

n da

mpi

ng

out t

he fr

ee v

ibra

tions

.

NASA/CP—2005-213655/VOL1 199

AT

IGA

TA

TIG

AT

3000

0 rp

m (

3141

rad

/sec

)

4000

0 rp

m (

4188

rad

/sec

)

RO

TO

R A

ND

FIN

GE

R R

ES

PO

NS

E T

O A

P

ER

IOD

IC F

OR

CIN

G F

UN

CT

ION

-0.0

012

-0.0

009

-0.0

006

-0.0

0030

0.00

03

0.00

06

0.00

09

0.00

12

00.

001

0.00

20.

003

0.00

40.

005

0.00

60.

007

Tim

e (

s)

Displacement (in)

Rot

orF

inge

r-10

psi

del

taF

inge

r-5

psi d

elta

Fin

ger-

1 ps

i del

ta-0

.001

2

-0.0

009

-0.0

006

-0.0

0030

0.00

03

0.00

06

0.00

09

0.00

12

00.

0005

0.00

10.

0015

0.00

20.

0025

0.00

30.

0035

Tim

e (

s)

Displacement (in)

Rot

orF

inge

r-10

psi

del

taF

inge

r-5

psi d

elta

Fin

ger-

1 ps

i del

ta

1000

0 rp

m (

1047

rad

/sec

)

2000

0 rp

m (

2094

rad

/sec

)

-0.0

012

-0.0

009

-0.0

006

-0.0

0030

0.00

03

0.00

06

0.00

09

0.00

12

00.

0005

0.00

10.

0015

0.00

20.

0025

Tim

e (

s)

Displacement (in)

Rot

orF

inge

r-10

psi

del

taF

inge

r-5

psi d

elta

Fin

ger-

1 ps

i del

ta-0

.001

2

-0.0

009

-0.0

006

-0.0

0030

0.0

003

0.0

006

0.0

009

0.0

012

00.

000

50

.001

0.0

015

0.0

02

Tim

e (

s)

Displacement (in)

Rot

orF

ing e

r-10

psi

del

taF

inge

r-5

psi d

elta

Fin

ger-

1 ps

i del

ta

The

rot

or p

ulls

aw

ay f

aste

r (n

atur

al f

requ

ency

of

the

fing

er

is lo

wer

than

the

forc

ing

func

tion

freq

uenc

y

Rot

or m

eets

fi

nger

1-ps

i fin

ger

free

fal

ls f

aste

r th

an th

e ro

tor,

but

has

to

follo

w th

e ro

tor

The

fre

e fa

ll o

f th

e 5-

psi

fing

er h

appe

ns a

t the

sam

e ra

te a

s th

e ro

tor

dow

n m

otio

n

Fin

ger

s d

o n

ot

follo

w t

he

roto

r to

o w

ell,

the

roto

r fr

equ

ency

bei

ng

mu

ch

hig

her

th

en t

he

nat

ura

l fr

equ

ency

of

the

fin

ger

s.

NASA/CP—2005-213655/VOL1 200

AT

IGA

TA

TIG

AT

-0.0

012

-0.0

009

-0.0

006

-0.0

0030

0.00

03

0.00

06

0.00

09

0.00

12

00.

0005

0.00

10.

0015

0.00

20.

0025

0.00

30.

0035

0.00

4

Tim

e (s

)

Displacement (in

Rot

or

FS

k=

11.7

2 lb

f/in

FS

k=

17.5

8 lb

f/in

FS

k=

23.4

4 lb

f/in

FS

k=

35.1

6 lb

f/in

FS

k=

36.8

8 lb

f/in

FS

k=

58.6

0 lb

f/in

FS

k=

70.3

2 lb

f/in

Ro

tor

and

Fin

ger

Res

po

nse

to

a P

erio

dic

Fo

rcin

g F

un

ctio

n

wit

h

freq

uen

cy

of

2094

ra

d/s

(2

0000

rp

m)

and

a

dro

p

thro

ug

h t

he

pre

ssu

re e

qu

aliz

atio

n h

ole

s o

f 5

psi

. F

ING

ER

S

TIF

FN

ES

S IS

VA

RIE

D

Flu

id s

tiff

nes

s

6852

lbf/

in

0.09

00.

090

70.3

2

0.07

50.

075

58.6

0

0.06

00.

060

46.8

8

0.04

50.

045

35.1

6

0.03

00.

030

23.4

4

0.01

50.

030

17.5

8

0.01

50.

015

11.7

2

LP

Fin

ger

Thi

ckne

ss

(in)

LP

Fin

ger

Thi

ckne

ss

(in)

K lb

f/in

Ks

that

allo

ws

smo

oth

er

roto

r/fi

ng

er

inte

ract

ion

NASA/CP—2005-213655/VOL1 201

AT

IGA

TA

TIG

AT

Wh

at d

oes

th

is m

ean

??

??

•Flu

id s

tiffn

ess

muc

h la

rger

than

that

of a

ny p

ract

ical

fing

er s

tiffn

ess

•The

fing

er fr

ee fa

ll po

rtio

n is

gre

atly

influ

ence

d by

sm

all c

hang

es in

eq

uiva

lent

fing

er s

tiffn

ess.

•The

tran

smis

sibi

lity

port

ion

(rot

or p

ositi

ve p

ush)

is n

ot a

ffect

ed in

the

prac

tical

ran

ge o

f fin

ger

stiff

ness

.

•A to

o th

in o

f a fi

nger

will

sta

y op

en (

up)

beca

use

it do

es n

ot p

osse

ss

enou

gh in

ner

sprin

g st

iffne

ss to

ove

rcom

e ev

en a

sm

all a

mou

nt o

fC

oulo

mb

fric

tion

•Con

vers

ely

a to

o th

ick

finge

r co

uld

caus

e ch

atte

r os

cilla

tion

•For

the

case

s st

udie

d he

re th

e be

st c

hoic

es s

how

ing

a fin

ger

pick

ing

up

smoo

thly

with

the

roto

r ar

e fo

r st

iffne

sses

bet

wee

n 35

and

58

lbf/i

n

fing

er

sn

mk

=ω

NASA/CP—2005-213655/VOL1 202

AT

IGA

TA

TIG

AT

So

me

no

tes

reg

ard

ing

pre

vio

us

two

slid

esS

om

e n

ote

s re

gar

din

g p

revi

ou

s tw

o s

lides

●In

fre

e vi

brat

ion

the

ener

gy o

f th

e fi

nger

ge

ts e

xpen

ded

even

tual

ly th

roug

h C

oulo

mb

fric

tion

●T

he f

ree

vibr

atio

n ho

wev

er d

oes

not c

arry

th

e da

y, b

ecau

se th

e fi

nger

re-

enco

unte

rs

the

roto

r●

The

nat

ural

fre

quen

cy o

f th

e fi

nger

rem

ains

th

e sa

me

NASA/CP—2005-213655/VOL1 203

AT

IGA

TA

TIG

AT

CO

NC

LU

SIO

NS

CO

NC

LU

SIO

NS

A

thre

e di

men

sion

al

Nav

ier-

Stok

es

base

d co

de

(CF

D-

AC

E+/

FE

MST

RE

SS)

was

ut

ilize

d to

an

alyz

e th

e th

erm

oflu

id b

ehav

ior

of a

mod

ifie

d F

S1 .

•T

he

pres

sure

pa

tter

ns,

mas

s fl

ows

and

load

car

ryin

g ca

pabi

litie

s of

thi

s st

ruct

ure

wer

e as

sess

ed.

It w

as f

ound

th

at e

ven

at a

low

er li

near

vel

ocit

y of

216

mps

(70

8 fp

s) t

he

geom

etry

pro

pose

d ha

s go

od li

ftin

g ca

pabi

lity.

.

NASA/CP—2005-213655/VOL1 204

AT

IGA

TA

TIG

AT

CO

NC

LU

SIO

NS

CO

NC

LU

SIO

NS

●T

he in

terp

lay

betw

een

the

rota

tion

indu

ced

pres

sure

ge

nera

tion

and

the

axi

al p

ress

ure

drop

s co

ntro

lled

by

the

HP

sid

e, is

dom

inat

ed b

y ro

tati

on a

t lo

w H

P s

ide

pres

sure

, but

it is

the

n ta

ken

over

by

the

axia

l pr

essu

re d

rop

whe

n th

e la

tter

bec

omes

larg

er t

hen

173k

Pa

at 2

16m

ps.

●T

he p

ress

ure

patt

erns

gen

erat

ed b

y th

is g

eom

etry

at

low

pre

ssur

e dr

ops

prov

e th

at t

he s

eal b

ehav

es in

the

fa

shio

n of

a m

ini-

slid

er b

eari

ng.

NASA/CP—2005-213655/VOL1 205

AT

IGA

TA

TIG

AT

CO

NC

LU

SIO

NS

(co

nt

CO

NC

LU

SIO

NS

(co

nt ’’

d)

d)

.

The

dyn

amic

mod

el i

ntro

duce

d a

sim

plif

ied

spri

ng-

mas

s-da

mpe

r eq

uiva

lent

to

th

e co

mpl

icat

ed

stru

ctur

e pr

esen

ted

by t

he F

S.

The

1-D

OF

num

eric

al e

xper

imen

ts c

once

ntra

ted

on

the

dete

rmin

atio

n of

the

pha

se s

hift

and

dis

plac

emen

t tr

ansm

issi

bilit

y Υ.

The

se t

wo

para

met

ers

indi

cate

how

wel

l an

d un

der

wha

t co

ndit

ions

the

fin

ger

will

fol

low

the

rot

or.

NASA/CP—2005-213655/VOL1 206

AT

IGA

TA

TIG

AT

CO

NC

LU

SIO

NS

CO

NC

LU

SIO

NS

●It

was

fou

nd t

hat

●(i

) th

e ph

ase

shif

t va

lues

incr

ease

d w

hen

flui

d st

iffn

ess

was

low

and

com

para

ble

to t

hat

of t

he s

tick

, ●

(ii)

the

pha

se s

hift

val

ue d

ecre

ases

wit

h fl

uid

dam

ping

incr

ease

,●

(iii)

the

com

bina

tion

of

smal

l kSE

qu a

nd la

rge

kFE

qu a

mou

nts

to a

tra

nsm

issi

bilit

y Υ

=1,a

nd a

rev

ersa

l in

rol

e le

ads

to v

ery

smal

l Υs,

●

(iv)

for

dam

ping

val

ues

low

er t

han

175

N.s

/m2

(1

lbf.

s/in

2) d

ampi

ng h

as n

o ef

fect

on Υ

and

NASA/CP—2005-213655/VOL1 207

AT

IGA

TA

TIG

AT

CO

NC

LU

SIO

NS

CO

NC

LU

SIO

NS

●(v

) ce

rtai

n co

mbi

nati

ons

of m

ass

and

flui

d an

d so

lid

stif

fnes

s le

ad t

o ve

ry la

rge Υ

whe

n th

e ro

tor

spee

d ap

proa

ches

the

nat

ural

fre

quen

cy o

f th

e sy

stem

●(v

i) in

the

2-D

OF

mod

el s

mal

l cha

nges

in K

stic

k ca

n si

gnif

ican

tly

affe

ct t

he r

otor

/sti

ck in

terp

lay.

NASA/CP—2005-213655/VOL1 208

ROLE OF DISTRIBUTED INTER-BRISTLE FRICTION FORCE ON BRUSH SEAL HYSTERESIS

Helen Zhao and Robert Stango

Marquette University Milwaukee, Wisconsin

Role of Distributed Inter-bristle Friction Force On Brush Seal

Hysteresis

By

Helen Zhao and Robert Stango

Department of Mechanical and Industrial EngineeringMarquette University, Milwaukee, WI

Email: [email protected] ; [email protected]

NASA/CP—2005-213655/VOL1 209

Intr

oduc

tion

and

Back

grou

nd

θ

θ

α

μ ξ

Δ*o

ω

Figu

re 1

•Int

erfe

renc

e pa

ram

eter

•In

war

d ra

dial

flo

w-in

duce

d lo

ad

q o •Con

tact

for

ce

F res

gene

rate

d at

in

terf

ace

of f

iber

tip

and

rot

or

•Loc

al o

ncom

ing

flow

of

gas

tow

ard

bris

tle

pack

qx

Figu

re 1

Bru

sh s

eal w

ith v

ario

us w

orki

ng lo

ads

* oΔ

NASA/CP—2005-213655/VOL1 210

Inte

r-br

istle

fric

tion

forc

e m

odel

Fron

t Pla

te

Bris

tle P

ack

Z

X

AA

t/2t/2

ψ

Bac

kP

late(a)

Dow

nstre

am

Ups

tream

(b)

g c z

X

ς

Sec

tion

A-A1

23

y

Figu

re 2

(a)D

epic

tion

of p

artia

l br

ush

seal

with

fro

nt

and

back

pla

te t

hat

cons

trai

n br

istle

pac

k

(b)S

ectio

n A-

A vi

ew,

depi

ctin

g th

e co

mpa

ctiv

elo

ad g

car

ound

bris

tle p

ack.

Th

e in

tera

ctiv

e fo

rces

of

thr

ee f

iber

s (1

, 2,

3)

are

stu

died

for

hyst

eres

is

phen

omen

on

NASA/CP—2005-213655/VOL1 211

Inte

r-br

istle

Fric

tion

Mod

el (

cont

’d)

Figu

re 3

•Th

ree

un-d

efor

med

ne

ighb

orin

g fib

ers

subj

ecte

d to

the

co

mpa

ctiv

elo

ad g

c

•D

efor

mat

ion

of

fiber

s un

der

com

pact

ive

load

gc.

NASA/CP—2005-213655/VOL1 212

Inte

r-br

istle

Fric

tion

Mod

el (

cont

’d)

Figu

re 4

(a)S

egm

ent

of t

he d

efor

med

fib

er s

ubje

cted

to

the

unifo

rm c

ompa

ctiv

elo

ad

g can

d tr

actio

n fo

rce

f A-A

'an

d f B

'-B

(b)s

impl

ified

mod

el

depi

ctin

g in

tera

ctio

n be

twee

n ne

ighb

orin

g br

istle

s as

uni

form

ly

dist

ribut

ed m

omen

t m

* al

ong

defo

rmed

fib

er

NASA/CP—2005-213655/VOL1 213

Inte

r-br

istle

Fric

tion

Mod

el--

-der

ivat

ion

of m

*

dNdf

dNdf

BB

AA

μμ

==

−′′

−;

cdN

dsg

=⋅

dsg

df

dsg

df

cB

B

cA

A

μμ==

−′

′−

dsr

gdm

wc

μ 2=

Lr

gm

wc

μ 2=

wcr

gm

μ2*

=If

Hex

agao

nalc

lose

d-pa

ck, t

hen

)60

cos

21(

2*

+=

wcr

gm

μ

Ref

er t

o Fi

g.3

and

Fig.

4:

1. D

iffer

entia

l fric

tiona

l for

ce d

f A-A

’an

d df

B’-B

2. D

iffer

entia

l mom

ent

dm:

3. R

esis

ting

bend

ing

mom

ent

m

4. D

istr

ibut

ed b

endi

ng m

omen

t pe

r un

it le

ngth

m

NASA/CP—2005-213655/VOL1 214

Gov

erni

ng E

quat

ion

Acco

rdin

g to

Eul

er-B

erno

ulli

Law

:

*re

sF

mE

IM

Mκ

=+

d dsφκ

=w

ith

Gov

erni

ng E

quat

ion:

2*

2co

s()

res

dE

Im

Fds

φα

μφ

=−

−−

Non

-dim

ensi

onal

for

m o

f go

vern

ing

equa

tion:

)co

s(2

2*

**

2*2

φμ

αφ

−−

−=

EIH

F

EIH

m

dsdre

s

NASA/CP—2005-213655/VOL1 215

Boun

dary

con

ditio

ns a

nd

cons

trai

nt c

ondi

tions

Boun

dary

con

ditio

ns:

1. s

lope

con

stra

int

at t

he b

ristle

orig

in, i

.e.

Φ=

0 at

s=

0

2. F

ree

of m

omen

t at

the

bris

tle t

ip, i

.e.

dΦ/d

s=0

at s

=L

Cons

trai

nt c

ondi

tions

:

;t

tx

xy

yξ

ξε

ε−

<−

<

00

cos

;si

nL

L

tt

xds

yds

φφ

==

∫∫

θξ

θ

ξθ

θ

ξξ

sin

)(

)si

n(

)co

s(co

s)

(

**

**

os

ss

ss

os

HR

RR

y

RR

HR

x

Δ−

+−

+=

+−

Δ−

+=

Whe

re,

NASA/CP—2005-213655/VOL1 216

Ecce

ntric

ity o

f Sh

aft

Figu

re 5

Ecce

ntric

m

ovem

ent

of

roto

r an

d di

spla

cem

ent

of

bris

tle p

ack

NASA/CP—2005-213655/VOL1 217

Num

eric

al R

esul

ts a

nd D

iscu

ssio

n

00.

020.

040.

060.

080.

10.

120.

14−

0.2

−0.

15

−0.

1

−0.

050

0.050.1

0.150.2

0.250.3

Δ* /H*

Fres H*2

/EI

FAH

*2

FBH

*2

FCH

*2

Δ g/H*

EI

EI E

I

Act

ual p

ath

for

unlo

adin

g cy

cle

Fic

titio

us p

ath

for

unlo

adin

g cy

cle

A

B

C

EIH

F res

2 *

Figu

re 6

Rel

atio

nshi

p be

twee

n

and

durin

g lo

adin

g an

d un

load

ing Δ*

for

a tr

ansi

tion

seal

with

R s

/H* =

8.9,

θ=

450

and

=

=

0.13

5.

Δ g/H

* sh

ows

the

posi

tion

whe

re f

iber

s ar

e “s

tuck

”,

i.e.,

cann

ot c

ompl

etel

y re

cove

r fr

om b

endi

ng

durin

g un

load

ing Δ***

HΔ

2*

*m

H EI

m

NASA/CP—2005-213655/VOL1 218

Num

eric

al R

esul

ts a

nd

Dis

cuss

ion

(Con

t’d)

00.

020.

040.

060.

080.

10.

120.

14−

0.2

−0.

15

−0.

1

−0.

050

0.050.1

0.150.2

0.250.3

Δ* /H*

Fres H*2

/EI

m1=

0m

2=0.

045

m3=

0.09

0m

4=0.

135

Δ gΔ g

Δ g H

*m

2 H

*m

3 H

*m

4

00.

020.

040.

060.

080.

10.

120.

140

0.02

0.04

0.06

0.080.1

0.12

0.14

m* H

*2/E

I

Δg/H*

Figu

re 7

(a)

Rel

atio

nshi

p be

twee

n di

men

sion

less

con

tact

for

ce a

nd

dim

ensi

onle

ss p

enet

ratio

n de

pth

for

=0,

0.0

45, 0

.090

, 0.1

35.(

Resu

lts

show

n ar

e fo

r R s

/H* =

8.9,

θ=

45o ;

(b)

rel

atio

nshi

p be

twee

n Δ g

/H*

and

non-

dim

ensi

onal

ben

ding

mom

ent

m

m

NASA/CP—2005-213655/VOL1 219

Num

eric

al R

esul

ts a

nd

Dis

cuss

ion

(Con

t’d)

00.

020.

040.

060.

080.

10.

120.

140

0.2

0.4

0.6

0.81

1.2

1.4

Δ* /H*

Fres H*2

/EI

θ 1=15

o

θ 2=30

o

θ 3=45

o

Δ g

Δ g

Δ g H*

θ 1

H*

θ 2

H*

θ 3 15

2025

3035

4045

0

0.02

0.04

0.06

0.080.1

0.12

0.14

θ (d

egre

e)

Δg/H*

Figu

re 8

(a)

Rel

atio

nshi

p be

twee

n di

men

sion

less

con

tact

for

ce a

nd

dim

ensi

onle

ss p

enet

ratio

n de

pth

for θ=

15, 30

and

45

degr

ees.

(R

esul

ts s

how

n ar

e fo

r a

bris

tle w

ith R

s/H

* =8.

9, Δ

* o=0,

=0.

135)

an

d (b

) Re

latio

nshi

p be

twee

n Δ g

/H*

and

bris

tle la

y an

gle θ

m

NASA/CP—2005-213655/VOL1 220

Conc

lusi

ons/

Sum

mar

yTh

e m

icro

-mom

ent

can

give

ris

e to

a d

elay

ed

filam

ent

disp

lace

men

t as

the

sha

ft u

nder

goes

tr

ansi

ent

excu

rsio

n an

d m

oves

rad

ially

tow

ard

bris

tle p

ack

(upl

oadi

ng).

H

owev

er, a

s th

e sh

aft

retu

rns

back

to

its c

once

ntric

po

sitio

n (d

ownl

oadi

ng),

the

fila

men

t CA

NN

OT

com

plet

ely

reco

ver

from

its

defo

rmed

pos

ition

and

re

mai

ns lo

cked

in a

n al

tern

ate

conf

igur

atio

n.

Cons

eque

ntly

, an

annu

lar

gap

is g

ener

ated

be

twee

n th

e fib

er t

ips

and

shaf

t su

rfac

e, w

hich

pr

omot

es b

rush

sea

l lea

kage

and

red

uces

tur

bo-

mac

hine

ry p

erfo

rman

ce.

NASA/CP—2005-213655/VOL1 221

Conc

lusi

ons/

Sum

mar

y (c

ont’d

)

In g

ener

al, f

or a

giv

en b

rush

sea

l, th

e an

nula

r ga

p in

crea

ses

linea

rly a

s th

e m

icro

mom

ent

m*

is

incr

ease

d.

The

brus

h se

al h

avin

g a

shal

low

est

lay

angl

e (1

5o)

resu

lts in

the

sm

alle

st a

nnul

ar g

ap, i

ndic

atin

g th

at a

br

ush

seal

des

ign

with

sha

llow

lay

angl

e is

leas

t pr

one

to h

yste

resi

s ph

enom

enon

, and

can

lead

to

impr

oved

per

form

ance

.

NASA/CP—2005-213655/VOL1 222

DOE/PG&E LNG-TURBOEXPANDER SEAL AND BEARING RETROFIT

Donald E. Bently, Dean W. Mathis, and G. Richard Thomas Bently Pressurized Bearing Company

Minden, Nevada

DOE/PG&E LNG-Turboexpander

Seal and Bearing Retrofit

2004 NASA Seal / Secondary Air System Workshop9 November 2004

Donald E. Bently, P.EDean W. Mathis

G. Richard Thomas, P.E.

Bently Pressurized Bearing Company1711 Orbit WayMinden, NV USA

NASA/CP—2005-213655/VOL1 223

DOE/PG&E LNG-Turboexpander

Seal and Bearing Retrofit

2004 NASA Seal / Secondary Air System Workshop9 November 2004

Donald E. Bently, P.EDean W. Mathis

G. Richard Thomas, P.E.

Bently Pressurized Bearing Company1711 Orbit WayMinden, NV USA

The U.S. Department of Energy (DOE), the Idaho National Engineering and Environmental Laboratories (INEEL), and Pacific Gas & Electric (PG&E) have been involved in the development of new technologies to produce liquefied natural gas (LNG) for diverse consumer use at a PG&E facility in Sacramento, California. Their goal is to establish packaged LNG facilities at/near conventional high-pressure natural gas letdown stations where high-pressure (750 to 900 psig) pipeline gas is throttled to distribution pressures of 50 to 60 psi.

Typically, in the past, this throttling process was achieved via pressure regulation valves. The new technology utilizes a turboexpander as the pressure throttling device. The turboexpander consists of a single stage radial inflow expansion turbine wheel on end of the shaft and a centrifugal compressor wheel on the other end of the shaft, in a conventional, double overhung arrangement. The benefit of the new process is to recover the energy of expanding pipeline gas from 750 psig to 60 psig, while at the same time recompressing a portion of the pipeline gas.

Initially, the turboexpander was delivered with shaft contacting brush seals behind both wheels, 17-4 PH shaft material, and two combination radial/thrust magnetic bearings. INEEL and PG&E worked with the turboexpander OEM from early 2002 through March 2004, unsuccessfully trying to commission the turboexpander. Due to stability issues, the turboexpander could not operate above 55 to 60,000 rpm. Required design speed is 70,000 rpm.

In late March 2004, Bently Pressurized Bearing Company was asked to solve the stability issue. This was accomplished by modifying:

- the existing shaft seal design, replacing the originally supplied brush seals with labyrinth seals (reduction in both axial loading and destabilizing tangential forces)

- the existing shaft material to a titanium alloy (12 percent gain in stability margin) - the combination radial/thrust bearing design from magnetic bearings to pressurized gas bearings (increase in load

carrying capability and stiffness with significant improvement in damping). The pressurized gas radial/thrust bearings utilizing the high pressure pipeline gas (methane) as the pressure source (425 psi) for the bearings and the gas distribution system as the low-pressure sink (60 psi) for the bearings.

This presentation documents the:

- operating characteristics of the turboexpander as originally supplied from the OEM with brush seals, 17-4 PH shaft material, and magnetic bearings,

- the design modifications made to the turboexpander, - operating characteristics of the turboexpander with the modified labyrinth seals, titanium alloy shaft material, and

pressurized radial and thrust gas(methane) bearings.

The turboexpander now successfully and continually operates at 70,000 rpm.

NASA/CP—2005-213655/VOL1 224

Project Definition❏ The US Department of Energy (DOE), the Idaho National Engineering

and Environmental Laboratories (INEEL), and Pacific Gas & Electric (PG&E) have developed new technologies to produce liquefied natural gas (LNG) for diverse consumer use.

❏ The goal is to establish packaged LNG facilities at / near conventional high-pressure natural gas letdown stations where the high-pressure (750 psig) pipeline gas is throttled to distribution pressure of 50 to 60 psi.

❏ The turboexpander consists of an expansion turbine stage on end of a shaft and a centrifugal compressor stage on the other end of the shaft, in a double overhung arrangement.

❏ The benefit of the new process is to recover the energy of expanding pipeline gas from 750 psig to 60 psig, while at the same time re-compressing a portion of the pipeline gas.

NASA/CP—2005-213655/VOL1 225

Project Definition❏ Original turboexpander design:

■ Brush shaft seals - expander and compressor wheels■ Combination radial / thrust magnetic bearings■ 17-4 PH shaft material■ Large Keyphasor hole / crude balancing■ 85,000 rpm design speed

❏ 667 ft/sec (455 mph) journal surface velocity❏ 450 ft/sec ≈typical maximum for most industrial turbomachinery

❏ Turboexpander failed to operate satisfactorily during previous two years of field testing (Jan 2002 – March 2004).■ Unacceptable seal leakage and thrust loading

❏ 300 ft/sec surface velocity ■ Lack of both radial and thrust load carrying capability of the

magnetic bearings■ Magnetic bearing control system instability■ Inability of magnetic bearings to operate above 55-60,000 rpm

NASA/CP—2005-213655/VOL1 226

Cross section of original turboexpander rotor/bearing assembly

Compressor Expander

Overall cross section of turboexpander. Original configuration with magnetic radial/thrust bearings, rolling element catcher bearings, brush type shaft end seals.

NASA/CP—2005-213655/VOL1 227

Shaft Material: 17-4 PHShaft Mass: 6.35 lbm

7.06 lbm (including wheels)Length: 10.2 in

Cross section of original turboexpander rotor/bearing assembly

Compressor Expander

Overall cross section of turboexpander. Original configuration with magnetic radial/thrust bearings, rolling element catcher bearings, brush type shaft end seals. Note that original shaft material is 17-4 PH hardened steel.

NASA/CP—2005-213655/VOL1 228

Brush Seals

Seal Plate

Brush Seal

Original design shaft end seals – radial brush seal design

NASA/CP—2005-213655/VOL1 229

Brush Seals

Brush Seal

Original design shaft end seal – radial brush seal design. This particular seal is one of several that failed during operation – potentially due to surface velocities and diff pressure across the seal.

NASA/CP—2005-213655/VOL1 230

Revised turboexpander seal geometry

Original shaft end seal design replaced with multi tooth labyrinth seals. Laby seals also included axial steps or swirl brakes.

NASA/CP—2005-213655/VOL1 231

Revised turboexpander shaft geometry

Physical Vapor Deposition (PVD) is the act of vaporizing a solid material within a vacuum chamber. The temperature of the vaporized material is much higher than the inside the chamber, thus allowing the material to physically condense onto a substrate and form a coating. Coatings produced by the PVD process are hard and have high density.

Shaft (Ti Alloy) Mass = 3.67 lbm (4.38 lbm including aluminum wheels)Overall Length = 10.20 in

Coating material: WC/C (tungsten-carbon-carbide)Friction coefficient (dry steel): 0.10 – 0.20 Coating process: Physical Vapor Deposition (PVD)Coating color: black-greyCoating structure: lamellar

Redesigned shaft: Ti Alloy, internal magnetic once per turn speed reference, WC-C PVD coating

NASA/CP—2005-213655/VOL1 232

Cross section of rotor/bearing assembly

Expander Compressor

Shaft

Bearings

Gas exhaust Backers

Machine CasePressurized gas manifold

Cross section of turboexpander with replacement externally pressurized gas bearings, laby seals, and new shaft installed

NASA/CP—2005-213655/VOL1 233

Bearing Geometry and Features

Bearing Bearing and Backer

Left picture is the bearing itself

Right picture shows the bearing installed in the backer which was utilized to adopt the bearing geometry into the existing turboexpancer casing. Oval port w/O-Ring groove is gas inlet port. O-ring used to seal the backer to the casing.

NASA/CP—2005-213655/VOL1 234

410,000 lbf/in stiffness100 lbf *sec/in damping

Predicted damped response for the turboexpander retrofitted with labyrinths seals, Ti alloy shaft, and pressurized gas bearings

Predicted damped response, from finite element model, of the turboexpander rotor startup utilizing the new laby seals, new externally pressurized gas bearings, and redesigned Ti Alloy shaft.

NASA/CP—2005-213655/VOL1 235

Predicted damped response for the turboexpander retrofitted with labyrinths seals, Ti alloy shaft, and pressurized gas bearings

410,000 lbf/in stiffness100 lbf *sec/in damping

Predicted damped response, from finite element model, of the turboexpander rotor startup utilizing the new laby seals, new externally pressurized gas bearings, and redesigned Ti Alloy shaft.

Based on known residual unbalance in the impellers, vibration response was predicted to be 70-90 degrees out of phase with the expander end leading.

NASA/CP—2005-213655/VOL1 236

Root locus stability analysis of redesigned turboexpander rotor/bearing system

δ = π ( )-αω

“Log Dec” loses information by taking the ratio ofGrowth Rate to Natural Frequency:

δ = π ( )-αω

“Log Dec” loses information by taking the ratio ofGrowth Rate to Natural Frequency:

α-α

ωd

Predicted Stability analysis via Root Locus methodology – predicted stable to 137,000 rpm, discounting aerodynamic effects

NASA/CP—2005-213655/VOL1 237

Test Stand Bearing Design Verification - to 55,000 rpm on air

From Expander EndFrom Compressor End

Rotor run up in externally pressurized gas (air) bearings at Bently Pressurized Bearing Company test stand – to 55,000 rpm

NASA/CP—2005-213655/VOL1 238

Installed field instrumentation

Viewed from the Expander End

Radial ProbesRTD Leads Thrust Probe Lead

Field commissioning: expander end instrumentation

NASA/CP—2005-213655/VOL1 239

Installed field instrumentation

Viewed from the Compressor End

Radial Probes RTD LeadsThrust Probe Lead

Field commissioning: compressor end instrumentation

NASA/CP—2005-213655/VOL1 240

Field static thrust load testViewed from the Compressor End

250 Pound Static Load

Thrust Probe Lead

Viewed from the Expander End

Dial Indicator –Axial Movement

250 lb static thrust load test – radial and thrust bearings pressurized with 450 psig air

NASA/CP—2005-213655/VOL1 241

Transient Startup – 6 Oct 2004

NASA/CP—2005-213655/VOL1 242

1X startup data – Bode plot

NASA/CP—2005-213655/VOL1 243

Direct and 1X shaft relative orbit plots from expander and compressor end.

Note that expander 1X phase leads compressor 1X phase by 74 deg and that compressor response is 33% larger than expander end. Refer back to predicted response on slide 13. Very good agreement between predicted and actual response.

NASA/CP—2005-213655/VOL1 244

Transient Startup – 6 Oct 2004

Full spectrum from expander end xy shaft relative probes, shows forward and reverse vibration frequencies up to 74,200 rpm. Response is all forward 1X, circular, symmetric response as would be expected from the externally pressurized gas bearings.

NASA/CP—2005-213655/VOL1 245

Thank You

❏ Questions?

NASA/CP—2005-213655/VOL1 246

Brandan Robertson, NASA-JSC/ES5 281-483-3732, Houston, TX

Advanced Docking/Berthing System NASA Seal Workshop

GRC

November 9-10, 2004

ADVANCED DOCKING/BERTHING SYSTEM

Brandan Robertson National Aeronautics and Space Administration

Johnson Space Center Houston, Texas

NASA/CP—2005-213655/VOL1 247

Bra

ndan

Rob

erts

on, N

ASA

-JSC

/ES5

281

-483

-373

2, H

oust

on, T

X

Out

line

•B

ackg

roun

d

•Fu

ture

Pro

gram

Nee

ds

•E

xist

ing

Syst

ems

•St

atus

•A

dvan

ced

Doc

king

/Ber

thin

g Sy

stem

(A

DB

S)

Ove

rvie

w

•K

ey S

eal R

equi

rem

ents

•E

arly

Sea

l Dev

elop

men

t Wor

k

NASA/CP—2005-213655/VOL1 248

Bra

ndan

Rob

erts

on, N

ASA

-JSC

/ES5

281

-483

-373

2, H

oust

on, T

X

Bac

kgro

und

Ber

thin

gre

fers

to m

atin

g op

erat

ions

whe

re a

n in

activ

e m

odul

e/ve

hicl

e is

pla

ced

into

th

e m

atin

g in

terf

ace

usin

g a

Rem

ote

Man

ipul

ator

Sys

tem

-R

MS.

Doc

king

ref

ers

to m

atin

g op

erat

ions

whe

re a

n ac

tive

vehi

cle

flie

s in

to th

e m

atin

g in

terf

ace

unde

r its

ow

n po

wer

.

RM

S B

erth

ing

Dep

ictio

n

NASA/CP—2005-213655/VOL1 249

Bra

ndan

Rob

erts

on, N

ASA

-JSC

/ES5

281

-483

-373

2, H

oust

on, T

X

Futu

re N

eeds

Fut

ure

Mat

ing

Syst

em C

apab

ility

Req

uire

men

ts:

–A

sys

tem

abl

e to

sup

port

a v

arie

ty o

f m

issi

ons:

CT

V/C

EV

/CR

V, l

unar

gat

eway

, M

oon,

and

Mar

s–

Lig

htw

eigh

t, fa

ult t

oler

ant s

yste

m th

at b

lend

s w

ell i

nto

vehi

cle

OM

L (

aero

)–

Cap

able

of

auto

nom

ous

rend

ezvo

us &

doc

king

–

Ber

thin

g ca

pabl

e fo

r m

odul

ar a

ssem

bly

and

vehi

cle

swap

-out

–So

ftw

are

reco

nfig

urab

le f

or a

ran

ge o

f ve

hicl

es a

nd o

pera

tions

–Fa

st s

epar

atio

n fo

r ra

pid

rele

ase

–M

odul

ar f

or m

aint

enan

ce a

nd s

ervi

cing

–C

onst

ella

tion

saf

ety

& r

elia

bilit

y go

als

–A

dapt

able

to I

SS

–C

rew

and

larg

e ca

rgo

tran

sfer

–Po

wer

, dat

a, a

nd f

luid

tran

sfer

–V

ehic

le to

veh

icle

mat

ing

(CR

V-C

TV

-oth

ers)

req

uire

s an

drog

ynou

s in

terf

ace

NASA/CP—2005-213655/VOL1 250

Bra

ndan

Rob

erts

on, N

ASA

-JSC

/ES5

281

-483

-373

2, H

oust

on, T

X

And

rogy

nous

Per

iphe

ral

Doc

king

Sys

tem

(A

PA

S)W

eigh

t: ~9

50 lb

s (

660

lbs

AP

DA

-600

1 +

276

lbs

avio

nics

) (

hatc

h no

t inc

l.)

Max

OD

: 69”

dia

Hat

ch P

ass

Thr

ough

: 31.

38”

dia

Sour

ce: J

SC-2

6938

, “P

rocu

rem

ent S

peci

fica

tion

for

the

And

rogy

nous

Per

iphe

ral D

ocki

ng S

yste

m f

or th

e IS

S M

issi

ons

Pas

sive

Com

mon

Ber

thin

g M

echa

nism

(P

CB

M)

Wei

ght:

680

lbs2

(440

lbs

PC

BM

+ 2

40 lb

s ha

tch)

H

atch

Pas

s T

hrou

gh: 5

0”sq

uare

Max

OD

: 86.

3”di

aSo

urce

: SSP

410

04, P

art 1

, “C

omm

onB

erth

ing

Mec

hani

sm to

Pre

ssur

ized

E

lem

ents

IC

D”

& S

SP 4

1015

, Par

t 1, C

omm

onH

atch

& M

echa

nism

s T

o P

ress

uriz

ed E

lem

ents

IC

D

Rus

sian

Pro

beW

eigh

t: 70

0 lb

s(5

50 lb

s co

ne +

150

lbs

avio

nics

)M

ax O

D: 6

1”di

aH

atch

Pas

s T

hrou

gh: 3

1.5”

dia

(app

roxi

mat

e)

Sour

ce: E

nerg

ia

Adv

ance

d D

ocki

ng/B

erth

ing

Syst

em (

AD

BS)

1

Wei

ght:

est.

750

lbs

(inc

lude

s el

ectr

onic

s &

hat

ch)

Max

OD

: 54”

dia

Hat

ch P

ass

Thr

ough

: 31”

dia

Sour

ce: X

-38

Pro

gram

Gro

up

1 AD

BS

curr

entl

y un

der

deve

lopm

ent

2 Bul

khea

d ha

tch

ring

str

uctu

re n

ot in

clud

ed

Exi

stin

g Sy

stem

s

NASA/CP—2005-213655/VOL1 251

Bra

ndan

Rob

erts

on, N

ASA

-JSC

/ES5

281

-483

-373

2, H

oust

on, T

X

Lim

itat

ions

of

exis

ting

sys

tem

s:–

Do

not m

eet 2

-fau

lt to

lera

nt, t

ime-

crit

ical

rel

ease

req

uire

men

t for

cre

wed

veh

icle

s•A

PAS

for

Shut

tle

reli

es o

n 96

bol

t EV

A to

mee

t 2nd

faul

t tol

eran

ce•C

BM

pow

ered

bol

ts in

nom

inal

ops

are

not

tim

e cr

itic

al a

nd a

re s

ingl

e fa

ult t

oler

ant

–U

niqu

e ac

tive

& p

assi

ve h

alve

s: p

recl

udes

veh

icle

-to-

vehi

cle

mat

ing

–D

o no

t sup

port

aut

onom

ous

oper

atio

ns

•No

auto

mat

ic m

atin

g of

flu

id, p

ower

(A

PA

S d

oes

have

a s

ingl

e po

wer

/dat

a co

nnec

tor)

an

d fo

rced

air

um

bili

cals

•CB

M c

anno

t mat

e to

unm

anne

d ve

hicl

es–

Stan

dard

ISS

rac

ks c

anno

t pas

s th

roug

h ex

isti

ng d

ocki

ng p

orts

–S

igni

fica

nt v

eloc

itie

s re

quir

ed to

pro

vide

ali

gnm

ent &

cap

ture

for

ces

–C

rit-

1 op

erat

ions

sup

port

ed b

y in

tens

ive

trai

ning

& a

naly

sis

–H

igh

part

cou

nt /

mec

hani

cal c

ompl

exit

y w

ith

sing

le p

oint

fai

lure

s (r

elia

bili

ty a

nd f

ailu

re

tole

ranc

e pr

oble

ms)

–B

erth

ing

mec

hani

sms

do n

ot d

ock

and

dock

ing

mec

hani

sms

do n

ot b

erth

–R

ussi

an s

yste

ms

are

supp

lied

by

a fo

reig

n ve

ndor

wit

h su

bsta

ntia

l eco

nom

ic c

once

rns

•Pur

chas

e of

add

itio

nal u

nits

ban

ned

by I

ran

Mis

sile

Pro

life

rati

on S

anct

ions

Act

of

1997

•V

ery

lim

ited

acc

ess

to e

ngin

eeri

ng d

ata

–S

yste

ms

desi

gned

and

/or

cert

ifie

d fo

r ve

ry f

ew c

ycle

s an

d sh

ort e

xpos

ure

life

Exi

stin

g Sy

stem

s

NASA/CP—2005-213655/VOL1 252

Bra

ndan

Rob

erts

on, N

ASA

-JSC

/ES5

281

-483

-373

2, H

oust

on, T

X

Cur

rent

Sta

tus

Adv

ance

d M

atin

g Sy

stem

Dev

elop

men

t A

ctiv

itie

s–

Exp

lora

tion

Syst

ems

Tec

hnol

ogy

Mat

urat

ion

Prog

ram

has

sel

ecte

d ad

vanc

ed

mat

ing

syst

ems

for

cont

inua

tion

duri

ng th

e re

cent

IC

P ac

tivity

.

–JS

C h

as b

een

deve

lopi

ng a

n ad

vanc

ed m

atin

g sy

stem

(A

DB

S) s

ince

199

6.

•O

rigi

nall

y in

tend

ed f

or u

se o

n th

e X

-38

proj

ect

•D

esig

ned

to s

uppo

rt b

oth

bert

hing

and

doc

king

ope

rati

ons

and

to p

rovi

de f

utur

e m

issi

on a

rchi

tect

ure

flex

ibil

ity

(cro

ss-c

utti

ng te

chno

logy

)•

The

cur

rent

des

ign

base

line

is a

X-3

8-si

zed

risk

red

ucti

on u

nit (

RR

U)

Ful

ly a

ndro

gyno

us in

terf

ace

Use

s el

ectr

omag

nets

and

clo

sed-

loop

for

ce-f

eedb

ack

for

soft

-cap

ture

Min

imal

ly s

ized

for

cre

w tr

ansf

erF

ully

inte

grat

ed g

roun

d ba

sed

syst

em to

sho

w T

RL

4 m

atur

ity

•W

ork

on C

onst

ella

tion

sca

le s

yste

m to

beg

in a

s R

RU

mat

ures

–L

ong-

dura

tion

seal

tech

nolo

gy (

seal

-on-

seal

) ha

s be

en id

enti

fied

as

a cu

rren

t de

sign

gap

.

NASA/CP—2005-213655/VOL1 253

Bra

ndan

Rob

erts

on, N

ASA

-JSC

/ES5

281

-483

-373

2, H

oust

on, T

X

CA

D I

mag

e

AD

BS

Ove

rvie

w

A N

ext-

Gen

erat

ion

Mat

ing

Mec

hani

sm–

Des

igne

d sp

ecif

ical

ly to

take

adv

anta

ge o

f m

oder

n el

ectr

omec

hani

cal t

echn

olog

y

–In

corp

orat

es th

e le

sson

s le

arne

d an

d ex

peri

ence

s fr

om p

revi

ous/

curr

ent m

atin

g m

echa

nism

dev

elop

men

t and

use

–D

esen

siti

zes

mat

ing

mec

hani

sm o

pera

tion

s an

d pe

rfor

man

ce f

rom

oth

er v

ehic

le s

yste

ms

requ

irem

ents

–S

uppo

rts

both

doc

king

and

ber

thin

g op

erat

ions

–S

uppo

rts

auto

nom

ous

rend

ezvo

us &

mat

ing

–A

ligne

d w

ith N

AS

A S

trat

egic

Pla

n

Inte

ract

ive

Ove

rvie

w

NASA/CP—2005-213655/VOL1 254

Bra

ndan

Rob

erts

on, N

ASA

-JSC

/ES5

281

-483

-373

2, H

oust

on, T

X

Key

Sea

l Req

uire

men

ts

•Se

al-o

n-se

al in

terf

ace

•V

ery

low

leak

rat

e

•L

ong

life

–L

ong-

dura

tion

expo

sed

peri

ods

–L

ong-

dura

tion

mat

ed p

erio

ds

–D

eep-

spac

e en

viro

nmen

ts

–M

ay a

lso

be a

pot

entia

l for

hig

h m

ate/

dem

ate

cycl

e lif

e

•R

edun

danc

y

NASA/CP—2005-213655/VOL1 255

Bra

ndan

Rob

erts

on, N

ASA

-JSC

/ES5

281

-483

-373

2, H

oust

on, T

X

AD

BS

Sea

l Loc

atio

ns

Inte

rfac

e Se

al

Hat

ch S

eal

3X L

ockd

own

Act

uato

r M

anua

l Inp

ut S

eal

3X L

atch

Act

uato

r M

anua

l Inp

ut S

eal

3X S

epar

atio

n S

yste

m

Man

ual I

nput

Sea

l

21X

Ele

ctro

nics

C

onne

ctor

Sea

lV

ehic

le S

truc

tura

l In

terf

ace

Seal

Bot

tom

Rin

g E

xten

sion

Sea

l

NASA/CP—2005-213655/VOL1 256

Bra

ndan

Rob

erts

on, N

ASA

-JSC

/ES5

281

-483

-373

2, H

oust

on, T

X

To

pres

erve

the

fully

and

rogy

nous

des

ign

conc

ept t

he s

eal d

esig

nap

proa

ch b

asel

ined

w

as a

sea

l-on

-sea

l im

plem

enta

tion

sim

ilar

to th

e A

poll

o So

yuz

(AST

P) s

eals

.

Subs

cale

sea

l-on

-sea

l ela

stom

eric

dev

elop

men

t with

Par

ker

Inc.

•Qui

ck d

evel

opm

ent a

nd te

stin

g to

eva

luat

e se

al-o

n-se

al p

oten

tial

•2 c

ross

-sec

tion

s (f

lat t

op a

nd e

llip

tica

l) a

nd 2

dif

fere

nt d

urom

eter

sil

icon

mat

eria

ls•H

eliu

m le

ak te

stin

g an

d se

al lo

ad f

orce

test

ing

com

plet

ed in

Jul

y 20

01•A

dhes

ion

test

ing

is o

ngoi

ng

Tes

t res

ults

•Lea

k ra

tes

com

para

ble

to I

SS C

BM

sea

ls w

ith

offs

et o

f 0.

050

inch

es a

nd n

o ga

ppin

g (~

20

conf

igur

atio

ns te

sted

)•C

ompr

essi

on f

orce

test

ing

show

ed th

at “

flat

top”

slig

htly

hig

her

than

“el

lipt

ical

”fo

r th

e 70

du

rom

eter

at (

96 &

87

lb/i

n) a

nd f

or th

e 50

dur

omet

er a

t (46

& 4

2 lb

/in)

. R

esul

ts in

dica

ted

that

sea

l-on

-sea

l in

the

“acc

epta

ble”

rang

e fo

r us

e.

•Adh

esio

n te

st r

esul

ts p

endi

ng; s

erie

s of

“bu

ttons

”m

olde

d fr

om e

ach

mat

eria

l are

cur

rent

ly

mat

ed a

nd c

ompr

esse

d fo

r ev

entu

al s

epar

atio

n an

d in

spec

tion

at T

BD

reg

ular

inte

rval

s of

ti

me.

Ear

ly A

DB

S Se

al D

evel

opm

ent

NASA/CP—2005-213655/VOL1 257

Bra

ndan

Rob

erts

on, N

ASA

-JSC

/ES5

281

-483

-373

2, H

oust

on, T

X

1.00

”

26”

rad

Tun

nel F

lang

e

Tun

nel F

lang

e

Seal

Ret

aine

r

Seal

Ret

aine

r

Seal

-on-

met

al

C/L

Seal

ing

Plan

e

RR

U I

nter

face

Sea

l Con

cept

NASA/CP—2005-213655/VOL1 258

Bra

ndan

Rob

erts

on, N

ASA

-JSC

/ES5

281

-483

-373

2, H

oust

on, T

X

Con

clus

ions

•N

o el

asto

mer

ic s

eal-

on-s

eal s

how

sto

pper

s ye

t

Forw

ard

wor

k•

As

soon

as

fund

ing

pict

ure

clea

rs u

p, m

ove

forw

ard

wit

h a

full

dev

elop

men

t sea

l pur

chas

e fo

r th

e R

RU

•N

eed

to e

stab

lish

“the

”ba

seli

ne s

eal c

ross

-sec

tion

•O

ptim

ize

seal

to g

uara

ntee

opt

imal

sea

ling

: per

cent

of

fill

, squ

eeze

, cro

wn

prof

ile

and

heig

ht, i

f el

asto

mer

ic•

Est

abli

sh to

tal p

oten

tial

sea

l mis

mat

ch: m

isal

ignm

ent,

ther

mal

exp

ansi

on, f

lang

e de

flec

tion

•E

stab

lish

acc

epta

ble

seal

for

ce a

nd le

ak r

ate

•D

eter

min

e if

a s

ingl

e pi

ece

~54”

seal

/ret

aine

r co

nstr

ucti

on p

ossi

ble.

Par

ker

has

indi

cate

d th

ey d

o th

is n

ow in

a n

ewly

acq

uire

d fa

cili

ty.

•E

valu

ate

conc

epts

and

res

ults

for

ful

l-sc

ale

Con

stel

latio

n im

plem

enta

tion

•E

valu

ate

RR

U d

esig

n up

war

d sc

alin

g•

Are

met

alli

c se

als

a be

tter

sol

utio

n?

Ear

ly A

DB

S Se

al D

evel

opm

ent

NASA/CP—2005-213655/VOL1 259

Bra

ndan

Rob

erts

on, N

ASA

-JSC

/ES5

281

-483

-373

2, H

oust

on, T

X

Bac

kup

Slid

es

NASA/CP—2005-213655/VOL1 260

Bra

ndan

Rob

erts

on, N

ASA

-JSC

/ES5

281

-483

-373

2, H

oust

on, T

X

Exi

stin

g Sy

stem

Sea

ls

Apo

llo S

oyuz

Tes

t P

rogr

amD

ocki

ng S

yste

m I

nter

face

Sea

l Dia

gram

NASA/CP—2005-213655/VOL1 261

Page intentionally left blank

NASA Glenn Research Center

Evaluation of Ceramic Wafer Seals for Future Space Vehicle Applications

Mr. Patrick H. Dunlap, Jr.Dr. Bruce M. Steinetz

NASA Glenn Research Center, Cleveland, OH

Mr. Jeffrey J. DeMangeUniversity of Toledo, Toledo, OH

2004 NASA Seal/Secondary Air System WorkshopNovember 9-10, 2004

Mr. Patrick H. Dunlap, Jr.Dr. Bruce M. Steinetz

NASA Glenn Research Center, Cleveland, OH

Mr. Jeffrey J. DeMangeUniversity of Toledo, Toledo, OH

2004 NASA Seal/Secondary Air System WorkshopNovember 9-10, 2004

EVALUATION OF CERAMIC WAFER SEALS FOR FUTURE SPACE VEHICLE APPLICATIONS

Patrick H. Dunlap, Jr., and Bruce M. Steinetz


Cleveland, Ohio

Jeffrey J. DeMange University of Toledo

Toledo, Ohio

NASA/CP—2005-213655/VOL1 263


♦High temperature, dynamic seals required in future space vehicles for sealing:• Perimeters of movable ramps in hypersonic engines• Heatshields and thermal protection systems (TPS) for aerocapture

and atmospheric entry systems• Perimeters of movable control surfaces

Hypersonic engine seals

Introduction & Background

Seal location

Controlsurface seals

Heatshield seals

High temperature, dynamic structural seals are required in many different locations on future space vehicles. Seals are required in advanced hypersonic engines to seal the perimeters of movable engine ramps for efficient, safe operation in high heat flux environments at temperatures from 2000 to 2500 °F. High temperature seals are also required around heatshields and within thermal protection systems for aerocapture and atmospheric entry systems. If a winged vehicle is used as an entry system, seals would also be required around the perimeters of movable control surfaces on the vehicle.

NASA/CP—2005-213655/VOL1 264


♦ Design requirements are demanding:• Restrict flow of hot gases at extreme temperatures

– Propulsion system applications: Up to 2500 °F with high heat fluxes

– Airframe applications:• 2100 °F for tile-based TPS• 2600+ °F for CMC systems

• Seal against distorted/curved surfaces

• Limit loads against sealing surfaces

• Stay resilient for multiple load/heating cycles

• Resist scrubbing damage

♦ Existing seals do not meet requirements

Seal Challenges & Design Requirements

Goal: Develop sealing systems that meet these requirements and demonstrate performance in relevant environments

CMC control surface

Sealgap

The design requirements for these seal applications are very demanding. The seals must restrict the flow of hot gases and survive at very extreme temperatures. For propulsion system applications, temperatures can reach 2500 °F in the presence of very high heat fluxes. For airframe applications, such as TPS and control surface seals, peak temperatures depend on the type of materials used in the structures surrounding the seals. If insulating tiles are used around the seals, temperatures can reach about 2100 °F. However, if hot structure ceramic matrix composite (CMC) materials are used, seal temperatures can reach 2600 °F or even hotter.

In addition to surviving at extreme temperatures, the seals must seal against distorted, curved surfaces while not applying excessive loads to those surfaces. The seals must remain resilient for multiple load cycles and heating cycles in order to stay in contact with the opposing sealing surface. These seals are used in dynamic applications in which the sealing surfaces move during each mission. Because these surfaces are often rough, the seals must be able to survive scrubbing against them without a large change in leakage performance.

Existing seals do not meet these challenging requirements, so the Seal Team at NASA GRC is developing sealing systems that do meet these requirements and demonstrating their performance in relevant environments.

NASA/CP—2005-213655/VOL1 265


♦ Ceramic wafer seals originally developed during NASP program♦ Preload device behind wafers maintains contact with sealing surface♦ Current design:

• Material: monolithic silicon nitride (Honeywell AS800)• Size: 0.5 in. wide x 0.92 in. long x 0.125 in. thick

Seal Specimens

0.5 in.

0.5 in.

The seals examined in this study were ceramic wafer seals that were originally developed during the NASP program. They are composed of a series of thin ceramic wafers installed in a channel in a movable panel and preloaded from behind to keep them in contact with the opposing sealing surface. Preload devices that have been considered for this application include helical compression springs and canted coil springs.

Materials that were evaluated for the wafer seals during the NASP program included a cold-pressed and sintered aluminum oxide, a sintered alpha-phase silicon carbide, a hot-isostatically-pressed silicon nitride, and a cold-pressed and sintered silicon nitride. A detailed analytical comparison of all the materials that were considered ranked the advanced silicon nitride ceramics as the most promising material for future consideration. Given that those tests were performed in the late 1980’s, considerable improvements have been made since then to produce stronger and tougher ceramic materials. Because of these improvements and the high ranking of silicon nitride as a candidate wafer seal material, GRC selected silicon nitride as the best candidate for these seals. The wafers tested in the current study were made of monolithic silicon nitride (Honeywell AS800) and were 0.5-in. wide, 0.92-in. tall, and 0.125 in. thick. They had corner radii of 0.050 in.

NASA/CP—2005-213655/VOL1 266


♦ Goal: Provide ~0.1-in. stroke to keep seal in contact with sealing surface

♦ Silicon nitride compression springs• Commercially available

• Potential for high temperature use (2000+°F)

• Evaluated 2 designs:

– Standard: 0.815 in. high x 0.520 in. diam.

– Modified: 0.694 in. high x. 0.435 in. diam.

♦ Canted coil springs• Unique load vs. displacement curve provides

nearly constant force over large range

• Spring modeling and development to be discussed in other presentations

Seal Preload Device Designs

Large working deflection of canted coil spring

5% 35%

Canted Coil Spring

Canted coil spring

Silicon nitride compression

springs

The high temperature seal preload devices that are being developed and evaluated would be installed behind the seals to ensure sealing contact with the opposing sealing surfaces. The requirements for these devices are also quite challenging as they must operate in the same environment and temperature as the seals. For the sealing applications being considered, the goal is to have the preload devices provide a stroke of about 0.1 in. to keep the seals in contact with their opposing sealing surfaces for multiple load and heating cycles.

Two types of preload devices have been focused on for these applications. The first type of device was a silicon nitride compression spring produced commercially by NHK Spring Co., Ltd. Because they are made of silicon nitride, these springs have the potential to be used as high temperature (2000+ °F) seal preloading devices. Two different designs were tested for the current study: a standard spring and a modified design. The main physical difference between these designs was that the standard design was larger than the modified design.

Canted coil springs have also been evaluated as part of this effort because of their unique load versus displacement curve that provides a nearly constant force over a relatively large deflection range. The Seal Team’s efforts to develop and model these types of springs will be discussed in other presentations that follow this one.

NASA/CP—2005-213655/VOL1 267


♦ Measure load vs. linear compression, resiliency, and stiffness of:• Springs alone• Wafer seals on top of springs

♦ Multiple load cycles at room temperature, 1600, 2000, 2200, & 2500 °F♦ Silicon carbide test fixtures (test temperatures up to 3000 °F)

Hot Compression Test Fixture

Seal (4 in. long)

Seal holder

Loading rod connected to

actuator

Box furnace rated to 3000 °F

Load cell below furnace

A series of hot compression tests were performed on the wafer seals and silicon nitride compression springs using GRC’s hot compression test rig. Compression tests were performed inside a box furnace capable of 3000 °F using the test setup shown here. These tests were performed to determine the resiliency and stiffness of the springs by themselves and for the wafers on top of the springs. Load versus displacement (i.e., linear compression) data was also generated by these tests. Test specimens were installed into a holder that rested on a stationary base. A movable platen attached to the actuator was translated up and down to load and unload the test specimens. Test specimens were loaded for multiple load cycles at a variety of test temperatures including room temperature, 1600 °F, 2000 °F, 2200 °F, and 2500 °F.

NASA/CP—2005-213655/VOL1 268


0.0

0.5

1.0

1.5

2.0

2.5

3.0

3.5

0.000 0.020 0.040 0.060 0.080 0.100

Linear compression (in.)

Forc

e, lb

f

70 F - Cycle 10

2000 F - Cycle 10

2200 F - Cycle 10

2500 F - Cycle 1

2500 F - Cycle 10

Compression Test Results: Standard Silicon Nitride Compression Springs

♦ No permanent set at room temperature, 2000 °F, or 2200 °F after 10 cycles♦ Observations from 2500 °F testing:

• Spring exhibited 44% permanent set after 10 cycles at 2500 °F• Some load-bearing capability still exists• Need to optimize material and design for multiple use at 2500 °F

28 lbf/in.

46 lbf/in.

28 lbf/in.

13 lbf/in.

31 lbf/in.

Springs show promise as high temperature seal preload devices

Permanent set

70°F, Cycle 102000°F, Cycle 102200°F, Cycle 102500°F, Cycle 12500°F, Cycle 10

17 lbf/in.

This figure shows the results for the compression tests performed on the standard silicon nitride compression spring design. For the tests performed at room temperature, 2000 °F, and 2200 °F, there was no permanent set or relaxation observed after 10 load cycles. For clarity, the figure only shows the curves for cycle 10 of these tests because they were almost identical to the curves for all other load cycles.

The test at 2500 °F produced somewhat different results than the tests performed at lower temperatures. The figure shows load cycles 1 and 10 for this test. For both of these load cycles, the peak loads were lower than those for the other tests. After 10 load cycles at 2500 °F the spring exhibited 44% permanent set but still had some load-bearing capability.

The results of these tests indicate that the silicon nitride compression springs show promise as high temperature seal preload devices, but some work needs to be done to optimize this material and spring design for use at 2500 °F.

NASA/CP—2005-213655/VOL1 269


Compression Test Results: Modified Silicon Nitride Compression Springs

♦ No permanent set at any temperature even for wafer seals on top of springs

♦ Some hysteresis for wafers on top of springs; possibly due to friction between wafers and groove side walls

*Note: 4 springs tested under wafer seals; 1 spring used for other tests

0.0

0.5

1.0

1.5

2.0

2.5

0.000 0.010 0.020 0.030 0.040

Linear Compression, in.

Fo

rce,

lb

f

70°F2000°F2200°F70°F, Wafers+Springs1600°F, Wafers+Springs2000°F, Wafers+Springs

Test Results for 10th Load Cycle65 lbf/in.

58 lbf/in.60 lbf/in.

57 lbf/in.

64 lbf/in.

60 lbf/in.

This figure shows the results for the compression tests performed on the modified silicon nitride compression spring design. The modified springs were smaller than the standard springs, allowing them to be installed in a groove behind the wafer seals so that compression tests could be performed on the sealing system as a whole. In all of these tests there was no permanent set or relaxation observed after 10 load cycles at room temperature, 1600 °F, 2000 °F, or 2200 °F. For clarity, the figure only shows the curves for cycle 10 of each test because they were almost identical to the curves for all other load cycles. For all of the tests performed on the silicon nitride springs by themselves, there was very little hysteresis in their load vs. linear compression data. For the tests performed with seals on top of the springs, there was virtually no hysteresis for the room temperature tests but a small amount for the tests at 1600 °F and 2000 °F. It is possible that during the high temperature tests, there was some small amount of friction between the wafers and the side walls of the seal groove that caused this hysteresis as the wafers and springs were unloaded during each load cycle.

NASA/CP—2005-213655/VOL1 270


Dwell Tests on Silicon Nitride Compression Springs

♦ Observed creep after 1.5 hrs under compression:• At 2000 °F load dropped by 13.5%

• At 2200 °F load dropped by 34%

♦ Still residual load on springs after 1.5 hrs

Time, hr.

0.00 0.25 0.50 0.75 1.00 1.25 1.50

Lo

ad, l

bf

0.0

0.5

1.0

1.5

2.0

2.5

2000°F 2200°F

Test Results for Standard Spring Design

Dwell tests were also performed on the standard silicon nitride compression springs in which the springs were held under load for an extended period of time. After an hour and a half under compression at 2000 °F the load dropped off by 13.5%. At 2200 °F the load dropped by 34% after an hour and a half. Although theloads dropped, the springs were still able to bear a good amount of load in both cases after an hour and a half of loading at high temperatures. If these springs are used as preload devices in future seal applications, this drop in load will have to be accounted for.

NASA/CP—2005-213655/VOL1 271


♦ Measure seal frictional loads and wear rates

♦ Test parameters:• Tests at room temperature and 1600 °F

– Inconel 625 rub surfaces– Rub surface roughness before testing =

6 μin• Test performed at 2000 °F

– Silicon carbide rub surfaces– Rub surface roughness before testing =

29 μin• Two sets of 32 wafers preloaded by silicon

nitride compression springs

• Seal gap size = 0.125-0.135 in.• 1000 scrub cycles; 2000 in. of scrubbing

Hot Scrub Test Fixture

Wafer seals

Seal holder

Inconel 625 rub surfaces

Wafer seals Seal holder

Silicon carbide rub surfaces

The main test rig that was used for the compression tests was also used to perform scrub tests on the seals using the set of test fixtures shown here. Tests were performed at room temperature, 1600 °F, and 2000 °F to evaluate seal wear rates and frictional loads as the seals were scrubbed against Inconel 625 and silicon carbide rub surfaces. The seals were installed in grooves in two stationary seal holders on either side of a pair of movable rub surfaces. The rub surfaces were assembled in a holder that was connected through the upper load train to the actuator.

Inconel 625 rub surfaces were used for the tests performed at room temperature and 1600 °F, while silicon carbide rub surfaces were used for the test performed at 2000 °F. The Inconel 625 rub surfaces had an average surface roughness before testing of about 6 μin. The roughness of the silicon carbide rub surfaces before testing was 29 μin. The gaps between the rub surfaces and the seals were set by spacer shims in front of and behind the seal holders. Gap sizes of 0.125-0.135 in. were used for these tests.

Four silicon nitride compression springs (modified spring design) were installed in the bottom of each seal groove to keep the wafer seals preloaded against both rub surfaces. A load transfer element was placed on top of the springs to support the wafers and distribute the load from the springs. Thirty two wafers were installed into each seal holder to fill the 4-in.-long seal grooves. The amount of compression on the seals and springs (0.030 in.) was set through an interference fit between the seals and the rub surfaces resulting in a preload of about 2 lb per inch of seal.

During these tests, the seals were held in place in the holders while the rub surfaces were scrubbed up and down against them. The seals were subjected to 1000 scrub cycles at 1 Hz for a total scrub length of 2000 in. for each test. Frictional loads were measured by the load cell under the furnace below the test fixture base. Seal wear rates were determined by examining the condition of the seals before and after each test and by measuring seal weight changes and changes in flow rates.

NASA/CP—2005-213655/VOL1 272


Scrub Test Results

♦Room temperature test against Inconel 625• Frictional loads peaked at 15.5 lb• Friction coefficient reached ~1• Rub surfaces became rougher

(43 μin)

♦1600 °F test against Inconel 625• Loads reached 23.5 lb• Max friction coefficient = 1.5 • Final rub surface roughness =

34 μin

♦2000 °F test against silicon carbide• Loads peaked at 12.8 lb• Friction coefficient ~1 by end of test; surface roughness stayed ~28-29 μin• No load spike at start of test; seals did not stick to surface during furnace heatup

Peak Frictional Loads During Scrub Testing

0

5

10

15

20

25

30

0 200 400 600 800 1000

Scrub Cycle

Lo

ad, l

bf

70°F - IN 6251600°F - IN 6252000°F - SiC

Peak frictional loads for each of the scrub tests are presented in this figure. For the test performed at room temperature against Inconel 625 rub surfaces, the frictional loads started around 6 lbf at the beginning of the test and gradually rose as the test proceeded until they reached about 15.5 lbf by the end of the test. Based on these loads, the friction coefficient at the end of the test was about 1.0. Before this scrub test, the average surface roughness of the rub surfaces was about 6 μin. After the test, the surface roughness had risen to about 43 μin. This increase in surface roughness during testing likely contributed to the increase in frictional forces as the test proceeded.

For the test performed at 1600 °F against Inconel 625 rub surfaces, the load peaked initially before quickly dropping and then slowly rising again to about 23.5 lbf by the end of the test. The friction coefficient reached 1.5 by the end of this test, corresponding well with an increased surface roughness of 34 μin.

For the test at 2000 °F using silicon carbide rub surfaces, lower loads were recorded than for the tests performed at lower temperatures against Inconel 625 rub surfaces. There was no load spike at the start of the test, indicating that the seals did not stick to the silicon carbide rub surfaces during furnace heatup. By the end of the test, the load peaked at about 12.8 lbf. The friction coefficient for this test reached about 1.0, and the surface roughness of the silicon carbide rub surfaces remained at about 28-29 μin, the same roughness as at the start of the test.

NASA/CP—2005-213655/VOL1 273


Scrub Test Results

♦ Little if any damage to silicon nitride wafers during scrub testing• No chips in wafers• Weight of wafer stacks almost identical before and after testing

♦ Superficial burnishing on Inconel 625 rub surfaces after testing♦ Some debris on SiC rub surfaces after testing; believed due to

abrasion of oxide layer

Wafer seals after 1600 °F scrub test against Inconel 625

Wafer seals before 1600 °F scrub test against Inconel 625

Wafer seals before 2000 °F scrub test against silicon carbide

Wafer seals after 2000 °F scrub test against silicon carbide

After the scrub tests were completed, the seals and rub surfaces were inspected for signs of damage. These figures show what the seals looked like before and after scrubbing at 1600 °F against Inconel and before and after scrubbing at 2000 °F against silicon carbide. The seals showed little if any damage after testing. None of the wafers were chipped or broken during testing, and the total weight of the wafers before and after testing was almost identical. Superficial burnishing was observed on the Inconel rub surfaces after testing, but this was not deemed to be significant. A small amount of debris was observed on the silicon carbide rub surfaces after testing (also visible in the photo of the wafers after this test). It is believed that this debris is due to abrasion of the oxide layer that forms on the silicon carbide rub surfaces at high temperatures.

NASA/CP—2005-213655/VOL1 274


Flow Test Results

Metered air supply

Cover plate

Gap = 0.135 in.

Test seal

Seal cartridgePlenum

Springs

Flow fixture

Flow Rates Before and After Scrubbing

0.00

0.20

0.40

0.60

0.80

1.00

1.20

0 20 40 60 80 100

Delta P (psid)F

low

per

in

ch o

f se

al (

SC

FM

/in

) Before RT scrubbingAfter RT ScrubbingBefore 1600°F scrubbingAfter 1600°F scrubbingBefore 2000°F scrubbingAfter 2000°F scrubbing

♦ No change in flow rates after scrubbing at room temp., 1600°F, or 2000°F♦ Flow rates were ~32 times lower than those for best braided rope seals♦ Flow rates lower for wafers tested at 1600°F and 2000°F due to tighter

wafer height tolerance (0.0005 in. vs. 0.001 in. for RT test)

Flow test results for the wafer seals before and after each scrub test are presented here for a gap size of 0.135 in. These tests were performed with four silicon nitride springs installed behind the wafers to keep them preloaded against the cover plate. Flow rates for the wafers before and after scrubbing were almost identical in each case. This is consistent with the observation that the wafers were not damaged during the scrub test. These results are encouraging because they show that the seals are still effective at blocking flow even after 1000 scrub cycles at room temperature. Additionally, flow rates for the wafer seals were up to 32 times better than those for the best braided rope seals.

Flow rates for the wafers tested at 1600°F and 2000°F were lower than those for the test performed at room temperature. This was because a tighter wafer height tolerance was used for the high temperature tests (0.0005 in.) than what was used for the room temperature test (0.001 in.).

NASA/CP—2005-213655/VOL1 275


Flow Test Results

♦ Recorded “seal activating” pressure behind seals during flow tests before and after 2000 °F scrub test• Augments preload devices to keep seals in contact with sealing surface

• Pressure behind seals was ~93% of pressure differential across seals

0

10

20

30

40

50

60

70

80

90

100

0 20 40 60 80 100

Pressure differential across seal, psid

Pre

ssu

re b

ehin

d s

eal,

psi

g

For the flow tests that were performed on the wafers before and after the 2000 °F scrub test, a new pressure measurement was made that was not recorded in the previous flow tests. This pressure measurement allowed the pressure behind the wafers to be recorded as a function of the differential pressure across the seals. The figure on the right presents pressure data behind the wafer seals versus the pressure differential across the seals for the flow test performed on the wafers that were scrub tested at 2000 °F. This plot shows a linear relationship between these pressures such that the pressure behind the seals was equal to about 93% of the pressure differential across the seals. The pressure behind the seals serves as a “seal activating” pressure that augments the preload devices to further preload the seals against the sealing surface and keep them in contact with it. This seal activating pressure phenomenon for wafer seals was observed previously for wafer seals tested during the NASP program.

NASA/CP—2005-213655/VOL1 276


♦ Wafer seals performed well in room temperature, 1600 °F, and 2000 °F scrub tests• No chips in wafers or any other signs of damage

• No change in flow rates after scrub testing

♦ Wafer seal flow rates were ~32 times lower than those for best braided rope seals

♦ Silicon nitride compression springs continue to show promise as high temperature seal preload devices; need to optimize materialand design for multiple use at 2500 °F

♦ Future work:• Investigate other wafer shapes and sizes• Investigate seal + preloading device combinations that satisfy requirements

at higher temperatures

Summary

Based on the results of these tests, the following conclusions were made:

1. The silicon nitride wafer seals performed very well in scrub tests performed at room temperature, 1600 °F, and 2000 °F. No chips or any other signs of damage were observed on the wafers after testing, and there was no change in flow rates past the seals after scrub testing.

2. Flow rates for the wafer seals were up to 32 times better than those for the best braided rope seals.

3. Commercially available silicon nitride compression springs continue to show promise as high temperature seal preload devices, but the material and design for these springs needs to be optimized for multiple uses at 2500 °F.

More work needs to be done to investigate seal and preloading device combinations that ultimately satisfy all of the seal requirements. The authors plan to investigate other wafer shapes and sizes to see if those changes affect seal durability, frictional forces, and flow rates. Longer scrub tests will also be performed at higher temperatures to examine seal durability.

NASA/CP—2005-213655/VOL1 277

On the Development of a Unique On the Development of a Unique Arc Jet Test Apparatus for Arc Jet Test Apparatus for

Control Surface Seal EvaluationsControl Surface Seal Evaluations

Joshua R. Finkbeiner, Patrick H. Dunlap, and Bruce M. SteinetzJoshua R. Finkbeiner, Patrick H. Dunlap, and Bruce M. SteinetzNASA John H. Glenn Research Center, Cleveland, OHNASA John H. Glenn Research Center, Cleveland, OH

Malcolm Robbie, Gus Baker, Arthur Erker, and Joe AssionMalcolm Robbie, Gus Baker, Arthur Erker, and Joe AssionAnalex Corp., Cleveland, OHAnalex Corp., Cleveland, OH

Seal WorkshopSeal WorkshopNovember 10, 2004November 10, 2004

ON THE DEVELOPMENT OF A UNIQUE ARC JET TEST APPARATUS FOR CONTROL SURFACE SEAL EVALUATIONS

Joshua R. Finkbeiner, Patrick H. Dunlap, and Bruce M. Steinetz


Cleveland, Ohio

Malcolm Robbie, Gus Baker, Arthur Erker, and Joe Assion Analex Corporation

Cleveland, Ohio

NASA/CP—2005-213655/VOL1 279

OutlineOutline

Future seal requirementsFuture seal requirementsTest fixture design objectivesTest fixture design objectivesSimulation of reentry environmentSimulation of reentry environmentTest fixture design and capabilitiesTest fixture design and capabilities

Modular seal cartridgeModular seal cartridgeModular flapModular flapMetalMetal--toto--ceramic flap transmission systemceramic flap transmission systemInstrumentationInstrumentationAngular positioningAngular positioning

NASA/CP—2005-213655/VOL1 280

Future Vehicle Seal NeedsFuture Vehicle Seal Needs

Survive higher Survive higher temperaturestemperatures

Proximity to outer mold lineProximity to outer mold lineCMC hot structures CMC hot structures (2400(2400ººF) with high F) with high conductivityconductivity

Compatibility with Compatibility with advanced flap materialsadvanced flap materialsImproved resiliency to Improved resiliency to follow gap openingsfollow gap openings

ElevonRudder

Future reentry vehicles will require more advanced seals than the current generation of vehicles (i.e. Shuttle). The seals must be capable of surviving higher temperatures for two primary reasons. Next-generation vehicles will be smaller in size than the Shuttle, and as a consequence, the seals will be closer to the outer mold line of the vehicle. Additionally, the next generation vehicles will employ CMC hot structures in place of the tile-insulated aluminum control surfaces currently used in Shuttle. These materials will be exposed directly to temperatures of (anticipated) 2400 ºF. Combined with the high thermal conductivity of these materials, new seal designs will be exposed to more severe thermal environments.

The use of CMC materials also introduces a second challenge in the design of new control surface seals. The seal materials must be chosen with care to ensure that the seal does not stick to the control surface material.

Finally, the seal must be able to maintain resiliency at the elevated temperatures. The seal gap size can change due to thermal expansion and movement of the control surface during flight. The seal must be capable of responding to gap size changes in these conditions.

NASA/CP—2005-213655/VOL1 281

Seal Test RequirementsSeal Test RequirementsRequire environment Require environment representative of reentryrepresentative of reentry

NASA JSC Arcjet simulates NASA JSC Arcjet simulates reentry environmentreentry environment

Heat FluxHeat Flux

Seal temp 1800Seal temp 1800--24002400ººFF

Pressure drop 56Pressure drop 56--100 psf100 psf

Exposure time ~15 minExposure time ~15 min

A simulated reentry environment is ideal for testing next-generation reentry vehicle control surface seals. This includes a combination of the heat flux, temperature, pressure drop, and exposure time encountered during hypersonic flight.

The arcjet facility at NASA Johnson Space Center will provide the simulated reentry heat. Depending on the nozzle used in the facility, the arcjet will expose the seal to temperatures of 1800-2200 ºF with a pressure drop of 56-100 psf. The facility is actively cooled, allowing it to simulate the reentry environment for 15 minutes, a typical reentry time for Shuttle.

NASA/CP—2005-213655/VOL1 282

Seal Test Fixture ObjectivesSeal Test Fixture Objectives

Test control surface seals in JSC arcjet tunnelTest control surface seals in JSC arcjet tunnelExpose seals to reentry conditionsExpose seals to reentry conditions

Survive arcjet environmentSurvive arcjet environment

Movement of control surface during hot testMovement of control surface during hot test

Variability of test parametersVariability of test parametersSeal shapes, sizes, and materialsSeal shapes, sizes, and materials

Fixture position (angular, spacial)Fixture position (angular, spacial)

NASA/CP—2005-213655/VOL1 283

Fixture DesignFixture Design

Insulating Tile

Water-cooledCopper Sides

C/SiC Flap

Water-cooledMotor Housing

Water-cooledBrake Housing

Seal Gap

Height Adjustment

Sting Arm

The design of the advanced seal arcjet test fixture provides flexibility for testing diverse seal designs at the same time as it is capable of surviving the harsh arcjet environment. The leading edge of the fixture contains the seal specimen, instrumentation, and provides the primary structure of the test fixture. The underlying structure of the leading edge is formed from water-cooled oxygen-free high-conductivity (OFHC) copper and is covered with insulating tile panels designed to minimize heat transfer to the structure. The tiles are coated with a high emissivity coating to further reduce the heat loading on the interior structure.

The leading edge is mounted against a C/SiC flap. The leading edge and flap are connected via a set of motor and brake systems housed inside water-cooled OFHC copper enclosures. The motor and brake systems allow the flap to move while the arcjet test is active. This simulates loading on the seal during reentry conditions, including the effects of scrubbing a rough surface against the seal at high temperature.

The sting arm supports the test fixture in the arcjet facility and holds the fixture in the arcjet flow. The sting arm allows movement of the fixture in the direction of the arcjet flow, while a high adjustment mechanism allows adjustment of the position of the fixture perpendicular to the flow stream.

NASA/CP—2005-213655/VOL1 284

Test Fixture CapabilitiesTest Fixture CapabilitiesAbility to move flap Ability to move flap during testduring test

Immersion in arc jet Immersion in arc jet flowflow

Modular seal cartridgeModular seal cartridge

Modular flap designModular flap design

Modify angle of attack Modify angle of attack and yaw angleand yaw angle

The flap can be moved while the arcjet is active while a test is in progress. This allows recording of dynamic changes in the seal environment and permits recording the seal response to these dynamic changes.

The seal specimen is mounted in a modular cartridge. This cartridge allows diverse seal shapes and materials to be tested in the fixture while minimizing changes to the fixture. New seal designs can be incorporated into the fixture by manufacturing a new low-cost cartridge instead of requiring expensive design changes to the entire fixture.

The flap is also modular, allowing not only the standard C/SiC flap, but also other designs (such as a tile-insulated aluminum flap). The attachment mechanism of other flap designs need only conform to the interface transmission system of the fixture.

The fixture also permits changes to the angle of attack of the leading edge and the yaw angle.

NASA/CP—2005-213655/VOL1 285

Modular Seal CartridgeModular Seal CartridgeRapid swapRapid swap--out of out of candidate sealscandidate seals

Tailor instrumentation Tailor instrumentation locations to specific seal locations to specific seal designdesign

Secondary seals Secondary seals eliminate sneak flowseliminate sneak flows

Permit proper evaluation Permit proper evaluation of candidate sealsof candidate seals

Prohibit hot gas Prohibit hot gas ingestion into fixtureingestion into fixture

The modular seal cartridge allows inexpensive testing of diverse seal shapes, sizes, and materials as well as fast turn-around time to incorporate new seal designs. The cartridge is formed from a series of insulating tile blocks mounted to a metal backing plate. The seal (in this case, a stack of ceramic wafers) fits between the two blocks. A secondary rope seal fits around the cartridge, sealing the edges of the test seal and preventing sneak flows into the internals of the leading edge. Each test seal has its own seal cartridge, providing a low-cost solution to expanding the capabilities of the test fixture. Furthermore, the instrumentation locations can be altered for different seal designs by altering the instrument locations in the tile blocks.

The seal cartridge attaches to a removable instrumentation tray which slides in and out of the bottom of the fixture along a series of guiding rails. These rails constrain the instrumentation tray such that the seal is compressed uniformly against the flap.

NASA/CP—2005-213655/VOL1 286

InstrumentationInstrumentation

ThermocouplesThermocouplesLeading EdgeLeading Edge

Above coveAbove cove

Above sealAbove seal

Behind sealBehind seal

Pressure TransducersPressure TransducersAbove sealAbove seal

Behind sealBehind seal

Seal performance measurements include both pressures and temperatures.

Thermocouple passages are machined into the top tile cover and seal cartridge and provide temperature measurements above the seal cove, inside the seal cove both immediately above and below the seal, and downstream of the seal cove, all at several locations along the width of the fixture. Additionally, temperatures are monitored in sensitive locations inside the test fixture (e.g. the pressure transducers, motor and brake, etc.) to ensure that electronic components are not exposed to temperatures above their design specifications.

Pressure taps are also embedded in the top insulating tile and in the seal cartridge. These lines pass through the tile material and connect to one of two manifold blocks. Each manifold block contains machined passages leading from the pressure taps to threaded connectors. Small pressure transducers thread into these connectors. This system allows fast-response pressure measurements while at the same time protecting the pressure transducers from the harsh arcjet environment. The manifold block permits the pressure transducers to fit into the small space provided in the leading edge.

NASA/CP—2005-213655/VOL1 287

FlapFlapModular design permits Modular design permits testing of diverse testing of diverse materialsmaterials

MR&D/GE C/SiC flapMR&D/GE C/SiC flapAnticipated test Anticipated test temperature of 2400temperature of 2400ººFF

Surface roughnessSurface roughness

TileTile--insulated insulated aluminum flapaluminum flap

7.9 in

C/SiC flap received from MR&D

The flap is a modular design, permitting diverse flap materials to be tested in the advanced seal arcjet test fixture. As with the modular approach to seal designs, the modularized flap permits testing of various flap materials and finishes with candidate seal designs.

The flap shown here is a C/SiC flap provided by MR&D and GE Power Systems. The flap is expected to be exposed to temperatures as high as 2400 ºF. The C/SiC material represents future sealing challenges in an environment including a combination of high heat loading, high temperature, rough sealing surface, and scrubbing conditions.

The modular design of the flap permits the use of other flap designs such as a tile-insulated aluminum flap characteristic of Shuttle.

NASA/CP—2005-213655/VOL1 288

Flap Transmission SystemFlap Transmission SystemTransmission interfaces Transmission interfaces cooled metal to hot flapcooled metal to hot flap

Differences in thermal Differences in thermal growthgrowthInhibits heat transferInhibits heat transfer

Drive PlateDrive PlateBoundary between metal Boundary between metal and ceramicand ceramicNN22--cooled pins interface to cooled pins interface to hot ceramic flaphot ceramic flapLocates motor and reduces Locates motor and reduces heat transferheat transfer

Brake prevents unpowered Brake prevents unpowered movement of flapmovement of flap

Motor

N2-cooled Pins

Motor MountingBrackets

N2 CoolingHoles

The flap transmission system is capable of altering the flap angle while the fixture is in the arcjet flow stream. At the same time, the transmission system holds the flap despite the differences in thermal expansion between the hot C/SiC flap and the cooled metallic components. Furthermore, the transmission system inhibits heat transfer from the flap to the motor and brake, keeping the temperatures of these critical components below their design temperature limits. The transmission system also permits manual adjustment of the angle of attack of the leading edge.

The key to the transmission system is the nitrogen-cooled drive plate. Nitrogen is fed into the plate through a channel around the outer diameter of the disk. Three pairs of cooling holes blow nitrogen onto the motor/brake face and onto the inboard labyrinth seal. These holes can be orificed or plugged completely in order to supply more nitrogen to three Inconel drive pins. Each pin has an internal Inconel tube which supplies nitrogen directly to the internal tip of the pin. The nitrogen flows along the inside surface of the hot pin and exits through holes near the plate. The flow of nitrogen maintains the pin tips below 1500 ºF, even though the flap temperature is approximately 2400 ºF. After exiting the pin, the nitrogen maintains a positive purge pressure in the motor/brake housings, preventing the ingestion of hot arcjet gases. This effect is boosted by the inboard labyrinth seal.

Heat transfer from the flap to the inboard labyrinth seal is inhibited by a silicon nitride insulating disk.

NASA/CP—2005-213655/VOL1 289

Variable AnglesVariable AnglesFlap angleFlap angle

Variable from 0Variable from 0°° -- 30+30+°°

Increments of 1.5Increments of 1.5°°

Angle of attackAngle of attackAlter seal environmentAlter seal environment

Variable from 0Variable from 0°° -- 9090°°

Increments of 6Increments of 6°° (finer (finer possible)possible)

Yaw angleYaw angleAllow flow along seal gapAllow flow along seal gap

Rudder seals or special Rudder seals or special flight conditionsflight conditions

Yaw Angle

Angle of Attack

Flap Angle

Three test fixture angles are capable of being independently adjusted.

The flap angle can be adjusted while the test fixture is exposed to the arcjet flow stream. The angle can be adjusted between 0º and 30º, with small negative angles possible. The brake constrains the flap angle to increments of 1.5º.

The angle of attack of the leading edge can be manually adjusted between tests. This permits alteration of the seal environment by manipulating the shock structure near the seal cove. The angle of attack can be easily adjusted between 0ºand 90º in increments of 6º. Finer angular increments can be obtained with some disassembly and reassembly of the fixture.

The fixture yaw angle can also be manually adjusted. This adjustment permits some of the arcjet flow to move along the seal gap, exposing the seal to higher heat fluxes. This simulates control surface environments such as rudders as well as special flight conditions where flow moves along seal gaps.

NASA/CP—2005-213655/VOL1 290

SummarySummary

Future reentry vehicles pose greater sealing Future reentry vehicles pose greater sealing challenges than shuttlechallenges than shuttle

Seal proximity to outer mold lineSeal proximity to outer mold lineHot (2400Hot (2400°°F) CMC control surfacesF) CMC control surfaces

Control surface seal test fixture permits seal Control surface seal test fixture permits seal tests:tests:

Under reentryUnder reentry--like conditions (arc jet)like conditions (arc jet)With variable position/geometryWith variable position/geometryUsing several diverse seal conceptsUsing several diverse seal conceptsAgainst various flap materials and shapesAgainst various flap materials and shapes

NASA/CP—2005-213655/VOL1 291

ScheduleSchedule

Prepare final drawings: 1Q FY05Prepare final drawings: 1Q FY05

Complete wood model test: 1Q FY05Complete wood model test: 1Q FY05

Complete fixture fabrication: 3Q FY05Complete fixture fabrication: 3Q FY05

Perform seal tests: FY06Perform seal tests: FY06

Checkout tests may be conducted in late FY05

NASA/CP—2005-213655/VOL1 292

Wood Model TestsWood Model TestsEnsure that tunnel does Ensure that tunnel does not block due to size of not block due to size of modelmodel

Establish tunnel Establish tunnel operating settingsoperating settings

Check for hot spots on Check for hot spots on modelmodel

Perform critical fit Perform critical fit checkschecks

The advanced seal arcjet test fixture is anticipated to be one of the largest objects tested in JSC’s Arcjet facility. The large size of the fixture has introduced the possibility that the fixture could block the JSC Arcjet tunnel. In this situation, the fixture deflects flow into the arcjet vacuum chamber, increasing the chamber pressure. The higher chamber pressure prevents the arcjet stream from expanding, resulting in a high enthalpy gas stream tightly focused on a small area of the fixture. This situation would result in the destruction of the fixture. Fortunately, this situation is avoided by an automatic tunnel shutdown sequence.

To permit the successful completion of seal tests, the test fixture must not block the arcjet tunnel. The use of a larger tunnel nozzle can prevent the fixture from blocking the tunnel, although this comes with the trade-off of lower seal temperature and heat flux. To determine the correct nozzle, a wood model of the test fixture was constructed and shipped to JSC.

The wood model will be placed in the arcjet stream for a few seconds. The short time increment will be enough to determine whether the fixture size is too large for a particular nozzle. Additionally, the wood model will reveal hotspots. If these hotspots occur in places not forseen, the fixture cooling system can be modified to account for the higher heat transfer in these locations.

NASA/CP—2005-213655/VOL1 293

Test Fixture AnalysisTest Fixture AnalysisTemperature profileTemperature profile

Within limits for OFHC Within limits for OFHC coppercopper

Below annealing Below annealing temperaturetemperature

Heating rates are Heating rates are conservativeconservative

Stagnation heating over Stagnation heating over front face, housingsfront face, housings

Side wall heating ratesSide wall heating rates

This is the temperature profile of the fixture before an insulating tile panel was added to the front face of the fixture.

NASA/CP—2005-213655/VOL1 294

Test Fixture SurvivalTest Fixture SurvivalWater cooling positioned Water cooling positioned in high heating areasin high heating areas

Insulating tile cover Insulating tile cover minimizes cooling of minimizes cooling of flowflow

Measures to prevent hot Measures to prevent hot gas ingestiongas ingestion

Labyrinth sealsLabyrinth seals

Positive purge pressurePositive purge pressure

Secondary sealsSecondary seals

MotorHousing

BrakeHousing

The cooling system for the leading edge and motor/brake housings. This shows the water passages before an insulating tile was added to the front face of the leading edge.

NASA/CP—2005-213655/VOL1 295

Brazing Trials for Main HousingBrazing Trials for Main Housing

Brazing trials on OFHC copper test specimens Brazing trials on OFHC copper test specimens Motor and brake housingsMotor and brake housings

SidewallsSidewalls

Passed hydrotest at 750 psi for 10 min.Passed hydrotest at 750 psi for 10 min.

Test specimen for side walls

Braze joint

Test specimen for motor and brake housings

Braze joints

NASA/CP—2005-213655/VOL1 296

INVESTIGATIONS OF HIGH-TEMPERATURE KNITTED SPRING TUBES FOR STRUCTURAL SEAL APPLICATIONS

Shawn C. Taylor

Case Western Reserve University Cleveland, Ohio

Jeffrey J. DeMange

University of Toledo Toledo, Ohio

Patrick H. Dunlap, Jr., and Bruce M. Steinetz


Cleveland, Ohio

NASA/CP—2005-213655/VOL1 297

Control Surface Seal Design Challenges

• Prevent hot gas ingestion around actuated structures during atmospheric reentry

• Protect underlying temperature-intolerant structures• Survive high temperature exposure (1800 – 2200°F)

• Maintain resiliency through multiple load/heating cycles

• Limit loads against sealing surfaces• Resist abrasion damage

Control surface seals are used to fill the gap at the interface of static panels and actuated control surfaces such as rudders and elevons. These seals protect underlying temperature-intolerant structures such as mechanical actuators from the hot gases encountered during atmospheric reentry. The seals are typically installed in a compressed state that is 80% of the nominal seal diameter (20% compression). This allows the seal to accommodate expansions and contractions of the gap as panels shift relative to one another during flight.

To be effective, control surface seals must maintain enough resiliency to survive multiple compressive cycles at temperatures ranging from 1800-2200°F and remain in contact with both sealing surfaces. Since control surface seals often experience scrubbing across the sealing surfaces as the control surfaces are actuated they must also be capable of resisting abrasion damage. In addition, these seals must also limit the force that they apply to contacted surfaces, as some surface materials (e.g., Shuttle tiles) are easily damaged if compressive loads are excessive.

NASA/CP—2005-213655/VOL1 298

Baseline Seal Design on the X-38

Seal installed between the rudders and fins on the X-38 CRV

Control surface seals are typically installed in locations like the rudder/fin section shown above on the X-38 Crew Return Vehicle (CRV). The baseline control surface seal to be discussed in this presentation was selected as the primary seal installed in this location to protect the CRV’s rudder drive motors from hot gas ingestion during atmospheric reentry.

NASA/CP—2005-213655/VOL1 299

Baseline Seal Design on the Space Shuttle

Seal Locations• Landing gear doors• Payload bay door vents• Orbiter external tank

umbilical door

Different sized versions of the baseline control surface seal are used as thermal barriers on the Space Shuttle. Seals are located around the landing gear doors, the payload bay door vents, and the orbiter external tank umbilical door.

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Baseline Seal Components

• Inconel X-750 knitted wire spring tube– Primary resilient element

• Saffil core– Limits hot gas flow

• 2-layer Nextel 312 ceramic fabric sheath– Provides a uniform sealing surface– Acts as a thermal barrier– Prevents loss of Saffil batting from seal core

Saffil core

Spring tube

Nextel 312 sheath

The baseline control surface seal is comprised of three parts:

1. A knitted Inconel X-750 spring tube which is the primary resilient element of the seal

2. A Saffil core that limits the flow of hot gases through the seal

3. A 2-layer Nextel 312 ceramic fabric sheath which provides a uniform sealing surface, acts as a thermal barrier, and prevents the loss of Saffil from the core of the seal through the walls of the knitted spring tube

NASA/CP—2005-213655/VOL1 301

Addressing Seal Inadequacies

• Permanent set and loss of seal resiliency occur above 1200°F

• Poor performance due to creep and low yield strength of Inconel X-750

• Potential Resiliency Improvements– Material substitution – Knit geometry modifications

Permanent set

Baseline Seal

Goal: Enhance seal resiliency through spring tube design improvements

The baseline control surface seal has shown significant levels of permanent set in previous compression tests at elevated temperatures. The image above presents a comparison between an untested seal and a seal that has been compressed and heated to 1900°F in a tube furnace. The loss of resiliency in the tested seal is attributed to a decrease in strength of the Inconel alloy at the elevated test temperature. Further testing has shown that resiliency loss becomes evident at test temperatures as low as 1200°F. This temperature is noticeably close to the temperature at which the yield strength of the Inconel X-750 alloy begins to sharply decrease. Based on these observations, efforts to enhance seal resiliency by improving the performance of the spring tube have become the primary focus. To enhance spring tube resiliency, material substitutions and knit geometry modifications are being investigated.

NASA/CP—2005-213655/VOL1 302

Spring Tube Knit Geometry

Knit Parameters– Number of Wire Strands – Courses per Inch (CPI) – Needles (N)– Wire Diameter – Tube Diameter

The schematic above shows a representative geometry of a spring tube. A defined set of knit parameters is used describe the spring tube geometry as follows:

•Number of Wire Strands: The number of individual wires that are knitted in a parallel fashion to form the spring tube. The baseline design has three wire strands, where the illustration above shows only a single strand.

•Courses per Inch (CPI): The number of individual courses (loops) counted per inch along the length of a spring tube. The baseline design has 4.9 CPI.

•Needles (N): The number of individual loops counted in a single rotation around the spring tube circumference. The baseline spring tube has 10 needles.

•Wire Diameter: The diameter of the individual wires that are knitted to make the spring tube. The baseline wire diameter is 0.009”.

•Tube Diameter: The outer diameter of the spring tube. The baseline tube diameter is 0.560”.

Spring tube loop density (LD) was defined to facilitate an equivalent comparison of modified spring tube geometries. This value combines both CPI and needles into a single parameter which represents the number of loops per square inch.

NASA/CP—2005-213655/VOL1 303

Spring Tube Material Selection

• Inconel X-750 baseline material

• Material substitution requirements– High temperature yield strength– Resistance to creep deformation at elevated temperatures– Available in wire form

• Rene 41 chosen as X-750 replacement– Improved high temperature yield strength– Better resistance to creep deformation

The spring tube in the baseline control surface seal is fabricated from Inconel X-750 wire. Inconel X-750 is a precipitation hardenable nickel based superalloy. To improve high temperature spring tube performance, material substitutions were investigated. Replacement materials were screened on the basis of availability of wire form, high temperature yield strength, and resistance to creep deformation at elevated temperatures. Based on these criteria, Rene 41, another precipitation hardenable nickel based superalloy, was selected as a potential near term replacement for the baseline material. Rene 41 is commercially available in wire form, and it has high temperature creep and yield strength properties superior to those of Inconel X-750. Oxide dispersion strengthened (ODS) alloys including PM 2000 and Inconel MA754 have high temperature creep properties better than Rene 41 above approximately 1550°F; however, these materials are difficult to obtain in wire form.

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Spring Tube Material Selection

Creep Resistance• Rene 41 superior to X-750• ODS alloys best at high temp.

Yield Strength• Rene 41 superior up to 1900°F• ODS alloys better above 1900°F

The plots above present yield and creep rupture strength values at multiple temperatures for candidate materials. On the plot of yield strength vs. temperature, it can be seen that Rene 41 is superior to the other alloys up to approximately 1900°F. After that temperature, the ODS alloys maintain higher yield strengths. The plot of rupture strength vs. temperature shows a similar trend. Up to approximately 1550°F Rene 41 has the highest rupture strength, but once that temperature is exceeded, the ODS alloys have better strength. It is notable that Rene 41 has better strength properties than the baseline Inconel X-750 material at all temperatures. These figures also suggest that if the ODS alloys were obtainable in wire form for knitting, they could hold promise for improved spring tube performance at temperatures near 2000°F.

NASA/CP—2005-213655/VOL1 305

Compression Test Fixture

Seal

Seal holder

Silicon carbide test fixture facilitates tests up to 3000°F

General test conditions– 4 in. specimen length– Preload = 0.2 lbf (0.05lbf/in.)– Multiple load cycles to 20%

compression– 250 s. dwell– Various test temperatures

To evaluate the performance of spring tubes and seals, samples were tested using a state-of-the-art test rig at NASA Glenn Research Center. This rig facilitates high temperature compression testing at temperatures up to 3000°F. This is made possible by the SiC test fixture shown above that can withstand the elevated test temperatures. Spring tubes were tested in 4” lengths. A preload 0.2 lbf was used in the testing to define uniform contact between the spring tube and the loading platen. This was necessary because the samples are not visible during high temperature testing, as the furnace doors are closed. The spring tubes were subjected to multiple load cycles to a nominal 20% compression. At maximum deflection, a dwell period of 250 seconds was used to simulate the predicted time of maximum heating during atmospheric reentry. Multiple test temperatures were used to evaluate the performance of the spring tube samples.

NASA/CP—2005-213655/VOL1 306

Knit Geometry Screening Tests

• Test matrix– Inconel X-750 spring tube specimens

• Baseline, 3 strand, 28 loop density

• Modified, 1 strand, 34 loop density

• Modified, 1 strand, 64 loop density

– Heat treated and non-heat treated samples– Single tests (no replicates)

• Additional test conditions– Room temperature and 1500°F– Flat platen spring tube support– 10 compression cycles

A preliminary set of tests was conducted to evaluate the impact of knit geometry modifications on spring tube resiliency and load generation. The investigated geometry parameters included the number of wire strands and spring tube loop density. Single strand designs having loop densities of 34 and 64 loops per in.2

were compared to the three strand 28 LD baseline design. Both heat treated and non-heat treated samples were tested. Only single tests were run during the screening stage due to scheduling limitations and sample availability. These tests were conducted at both room temperature and 1500°F, and consisted of 10 loading cycles. Spring tubes were supported with a flat platen to simplify contact conditions allowing test results to be used in support of numerical modeling efforts.

NASA/CP—2005-213655/VOL1 307

Loop Density Effects at 1500°F

Loop density increase from 34 to 64• Improved cycle 10 resiliency by 21%• Had a negligible effect on cycle 10 peak load

Effects plots were used to highlight trends in the collected data. In the effects plot above, residual interference (seal spring-back) at the start of compression cycle 10 is plotted against loop density. A greater line slope suggests a more dominant influence of the factor on the response variable, which in this case is residual interference. A zero slope would indicate that the factor had no influence on the response. As shown above, increasing spring tube loop density from 34 to 64 improved post deflection spring back by approximately 21%. The effect of loop density on cycle 10 load was determined to be negligible.

NASA/CP—2005-213655/VOL1 308

Strand Effects at 1500°F

Increasing the number of strands from 1 to 3• More than doubled cycle 10 peak load• Had a negligible effect on cycle 10 resiliency

In the effects plot above, the influence of the number of wire strands on the peak load generated by the spring tube at maximum deflection is highlighted for testing at 1500°F. Increasing the number of strands from 1 to 3 more than doubled the cycle 10 peak load. The change in the number of wire strands had a negligible effect on the resiliency.

NASA/CP—2005-213655/VOL1 309

Material Selection Tests

• Test matrix– Inconel X-750 and Rene 41 spring tube specimens

• Baseline knit geometry

• Heat treated and non-heat treated samples

– Rene 41 tests were repeated once– Inconel X-750 tests were not replicated

• Additional test conditions– Room temperature, 1200°F, 1500°F, 1750°F, and 2000°F– Grooved seal holder spring tube support– 20 compression cycles

To evaluate the impact of material selection on spring tube performance, baseline geometry spring tubes fabricated from Inconel X-750 and Rene 41 were tested at room temperature, 1200°F, 1500°F, 1750°F, and 2000°F. Spring tubes of both materials were tested in the as-knitted (non-heat treated) and heat treated state. Material selection tests used a grooved seal holder to support the samples and were extended to 20 compression cycles instead of 10. These procedural changes were used to create a test that was more representative of typical re-entry conditions. Other test variables such as dwell time and compression level remained the same as in previous tests. Tests on Rene 41 samples were repeated once, whereas tests on the Inconel X-750 samples were not.

NASA/CP—2005-213655/VOL1 310

Rene 41 Performance Improvements

• Rene 41 spring tubes showed large resiliency improvements over Inconel X-750 specimens

– Maintained >95% resiliency through 1200°F

– Exhibited a 5.2x resiliency improvement at 1500°F after 20 cycles – Maintained reasonable resiliency up to 1750°F

The impact of material substitution (Rene 41 for Inconel X-750) is highlighted in the bar chart above, where average residual interference (spring back) at the start of cycles 2 and 20 is plotted against test temperature. Unlike the baseline Inconel X-750 spring tubes, Rene 41 samples maintained greater than 95% resiliency through 1200°F for all 20 compression cycles. At 1500°F, Rene 41 spring tubes showed a 5.2x improvement during cycle 20 over the baseline design. At 1750°F, Rene 41 spring tubes maintained a 28% cycle 20 residual interference when all resiliency was lost in the Inconel specimens. Despite the performance improvements achieved by the interim material substitution, further material substitutions will be necessary in order to satisfy the 1800°-2200°F design requirements, as all resiliency was lost by cycle 20 in both spring tubes at 2000°F.

NASA/CP—2005-213655/VOL1 311


Rene 41 substitution enhanced temperature capability ~275°F

The chart above plots cycle 20 residual interference vs. temperature for both spring tube materials to highlight the improvement in temperature capability attained through material selection only. For an arbitrary 75% spring back, the substitution of Rene 41 for Inconel X-750 showed a temperature improvement of ~275°F.

NASA/CP—2005-213655/VOL1 312


Pre-test 2000°F1750°F1500°F1200°F

20% Compression Tests on IN X-750 Spring Tube

Permanent set

20% Compression Tests on Rene 41 Spring Tube

Pre-test 1200°F 1500°F 1750°F 2000°F

Significant permanent set in Inconel at 1500°F

No visible permanent set in Rene until 1750°F

Calculated resiliency improvements resulting from material substitution were visually confirmed during post test inspection of the spring tube samples. As shown above, permanent set in the Inconel X-750 specimens was clearly evident at 1500°F. In the Rene 41 samples, however, resiliency loss was not visually distinguishable until 1750°F.

NASA/CP—2005-213655/VOL1 313


• Improved resistance to load relaxation during dwell at 1500°F – Force during cycle 1 dropped ~17% for Rene 41 vs. 40% for Inconel

• Produced higher peak loads than Inconel – 26% load increase at room temperature

– Loads still within the 5 lbf/in. limit for shuttle tiles

Rene 41 spring tubes knitted using the baseline knit geometry produced higher peak loads throughout all test cycles at 1500°F and resisted load relaxation during dwell better than the baseline Inconel X-750 spring tube samples. The plot of load vs. time above shows that during compression cycle 1, the load generated by the Inconel X-750 spring tube decreased approximately 40%, whereas the Rene 41spring tube suffered only a 17% load relaxation. Even though the Rene 41 spring tube generated room temperature peak loads that were approximately 26% higher than those produced by the Inconel X-750 spring tube, contact forces were still well below the 5 lbf/in. load limit for Space Shuttle tiles.

NASA/CP—2005-213655/VOL1 314

Heat Treatment Effects

• Heat treatment of Rene 41 spring tubes improved cycle 20 resiliency 9.3x at 1750°F

• Heat treatment did not improve Inconel spring tube performance

A comparison was also made during screening tests to determine the effects of heat treatment on the performance of the individual materials. As illustrated above, heat treated Rene 41 maintained much better high temperature resiliency than the non-heat treated samples. At the start of cycle 20 at 1750°F, heat treated Rene 41 spring tubes showed a 9.3x higher residual interference than the non-heat treated specimens. This behavior was not observed in tests of the baseline Inconel X-750 spring tubes.

NASA/CP—2005-213655/VOL1 315

Modification Effects on 1500°F Resiliency

• Material substitution improved cycle 10 resiliency by 2.5x• Rene 41 heat treatment improved resiliency by 93%• Loop density increase from 34 to 64 improved resiliency ~21%• Effect of number of strands was negligible

The effects plot above summarizes the relative influence of each of the evaluated spring tube factors on average residual interference at the start of cycle 10 for tests ran at 1500°F. As shown, the most significant factors were material selection and heat treatment of the Rene 41 spring tube. Material substitution of heat treated Rene 41 improved cycle 10 resiliency by approximately 2.5x when compared to the baseline Inconel X-750 spring tube. Heat treatment of the Rene 41 spring tube increased resiliency by 93% over the non-heat treated sample. Geometry effects were determined to be less significant; however, results did show that moderate performance improvements through knit geometry modification were feasible. Increasing the loop density in single stranded Inconel X-750 spring tubes from 34 to 64 improved residual interference by approximately 21%. Resiliency changes due to increasing the number of wire strands from 1 to 3 were determined to be negligible.

NASA/CP—2005-213655/VOL1 316

Further Improvements

• Rene 41 material substitution for Inconel X-750 yielded ~275°F temperature improvement

• Further improvements needed to reach design goal of 2200°F– Material selection– Heat treatment– Geometry

• Determination of dominant deformation mechanism– Identify creep or time-independent plastic flow as primary

cause of resiliency loss • Eliminated test dwell periods and increased load rates to

minimize ‘time effects’• Tested heat treated Rene 41 spring tubes (standard geometry)

at 1750°F

Screening tests showed that a material substitution of heat treated Rene 41 for the baseline Inconel X-750 alloy could increase temperature capabilities of the baseline knit geometry spring tube by ~275°F for an arbitrary residual interference of 75%. Preliminary geometry modifications also showed potential for enhancing spring tube performance. However, to reach the operating temperature design goal of 2200°F, further spring tube improvements will be needed. Utilizing data collected through the effects screening tests, material selection, heat treatment, and knit geometry will be revisited as opportunities for additional spring tube performance enhancement.

Previous tests showed a correlation between the temperature at which resiliency loss was detected and the temperature at which the alloys being tested were known to suffer significant decreases in yield strength. Permanent deformation in the spring tubes was suspected to be the result of both creep deformation and non time dependent plastic flow, but the dominant deformation mechanism was unknown. To distinguish between the two mechanisms, standard test conditions were modified to minimize the ‘time effects’ influencing the experimental results. Dwell periods were eliminated, and load rates were increased to limit the amount of creep deformation occurring during the tests. These tests were conducted at 1750°F using heat treated Rene 41 spring tubes with standard geometries and a 20% compression level.

NASA/CP—2005-213655/VOL1 317

Deformation Mechanism Identification

Eliminating dwell and increasing load rates at 1750°F• Showed higher resiliency: 2.1x at cycle 10 and 2.5x at cycle 20• Identified creep as dominant mechanism causing permanent set

The bar chart above plots average residual interference vs. compression cycle to illustrate the effect of eliminating the dwell periods and increasing load rate of the standard spring tube compression tests. Calculated resiliencies for spring tubes tested without dwell periods at maximum deflection were 2.1x higher at the start of cycle 10 and 2.5x higher at the start of cycle 20 when compared to the corresponding resiliencies calculated from standard compression tests with dwell periods. These results provided a strong indication that creep deformation was the dominant mechanism that will need to be addressed in order to further improve spring tube performance.

NASA/CP—2005-213655/VOL1 318

Spring Tube Microscopy

• Enhanced performance of larger grain, heat treated Rene 41 supports creep deformation theory

• Improved resiliency may be achieved through modified Rene 41 heat treatments or ODS alloy substitution

Heat Treated Inconel X-750

Non-Heat Treated Rene 41

Heat Treated Rene 41

To support the theory identifying creep as the dominant deformation mechanism leading to permanent set in tested spring tubes, a microstructure evaluation was conducted. Untested samples of heat treated Inconel X-750, non-heat treated Rene 41, and heat treated Rene 41 spring tubes were mounted in epoxy, sectioned, and examined using an optical microscope. As shown above, heat treated Inconel X-750 and non-heat treated Rene 41 samples had comparable grain sizes. These two spring tubes performed similarly during high temperature compression testing, taking on significant permanent set, whereas larger grain, heat treated Rene 41 spring tubes showed significant resiliency improvement. These findings support the theory of creep dominance, as large grain metals are classically better for resisting creep deformation than fine grain alloys. Based on this, improved resiliency may be achieved through modified Rene 41 heat treatments or ODS (oxide dispersion strengthened) alloy substitution that would provide an enhanced resistance to creep deformation.

NASA/CP—2005-213655/VOL1 319

DOE Analysis

• DOE based on screening tests• Results will be combined for an

optimized spring tube design

Heat Treatment• Rene 41 baseline• Rene 41 modified for

enhanced γ’ growth

Knit Geometry• Number of strands

• 3 strands

• 10 strands

• Wire diameter• 0.005” dia.

• 0.009” dia.

• Loop Density• 28 LD

• 127 LD

Spring tube samples ordered for testing in early December

Factors and Levels

In order to apply and test the concepts for improving spring tube performance generated through preliminary testing, a formalized design of experiments (DOE) study will be conducted. The factors to be tested include heat treatment, number of strands, wire diameter, and loop density. The levels for these factors are presented above. Spring tubes will be fabricated from Rene 41 and tested using previously defined test parameters, including a 20% nominal compression, 1750°F test temperature, and a 250 second dwell period at maximum deflection during each of the 20 compression cycles. Results from these tests will be used to produce an ‘optimized’ spring tube design which will then be incorporated into a control surface seal for evaluation.

NASA/CP—2005-213655/VOL1 320

Conclusions

• Current baseline control surface seal is inadequate for future space vehicles

• Heat treated Rene 41 material substitution for Inconel X-750 yielded ~275°F temperature improvement

• Material substitution showed resiliency improvements as high as 5.2x at cycle 20, 1500°F

• Preliminary geometry modification showed resiliency improvements as high as 21% at cycle 10, 1500°F

• Creep deformation identified as dominant mechanism leading to resiliency loss

• DOE analysis will be used to optimize Rene 41 spring tube design

• Further material substitution may be needed to reach 2200°F design goal

NASA/CP—2005-213655/VOL1 321

MODELING OF CANTED COIL SPRINGS AND KNITTED SPRING TUBES AS HIGH-TEMPERATURE SEAL PRELOAD DEVICES

Jay J. Oswald and Robert L. Mullen Case Western Reserve University

Cleveland, Ohio

Patrick H. Dunlap, Jr., and Bruce M. Steinetz National Aeronautics and Space Administration


NASA/CP—2005-213655/VOL1 323

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w p

ath

(C)

(A)

is s

tret

ched

(B)

is c

ompr

esse

d

NASA/CP—2005-213655/VOL1 328

Ver

ifica

tion

of S

prin

g B

ehav

ior

Mod

el

Com

pres

sion

Ten

sion

B

A

A

B

FE

A r

esul

ts s

how

com

pres

sion

and

tens

ion

as p

redi

cted

in b

ehav

ior

mod

el

Can

ted

co

il sp

rin

gs

sto

re e

ner

gy

in b

end

ing

rat

her

th

an t

ors

ion

NASA/CP—2005-213655/VOL1 329

Com

paris

on o

f Exp

erim

enta

l and

Ana

lytic

al

Res

ults

Obs

erva

tions

–G

ood

agre

emen

t in

beha

vior

and

mag

nitu

de

–M

odel

acc

urac

y de

crea

ses

with

hig

her

stiff

ness

spr

ings

–M

odel

initi

ally

und

eres

timat

es s

prin

g st

iffne

ss

NASA/CP—2005-213655/VOL1 330

Effe

ct o

f Spr

ing

Leng

th

•S

prin

g st

iffne

ss

depe

nds

on le

ngth

–S

hort

er s

egm

ents

ar

e so

fter,

mor

e no

n-lin

ear

–S

hort

er s

prin

gs

com

pres

s ax

ially

un

der

tran

sver

se

load

ing

NASA/CP—2005-213655/VOL1 331

Effe

ct o

f Fric

tion

on C

ante

d C

oil S

prin

gs

•F

rictio

n is

mod

eled

w

ith a

Cou

lom

b ap

prox

imat

ion

(F =

μN

)

•In

crea

sing

fric

tion

incr

ease

s th

e st

iffne

ss o

f the

sp

ring

NASA/CP—2005-213655/VOL1 332

Sum

mar

y of

a C

ante

d C

oil S

prin

g P

aram

eter

Stu

dyP

aram

eter

stu

dy–

wid

th–

ecce

ntric

ity–

coils

per

inch

–w

ire d

iam

eter

–ca

nt a

mpl

itude

Incr

easi

ng c

ant a

mpl

itude

➠

Incr

easi

ng w

ire d

iam

eter

➠

NASA/CP—2005-213655/VOL1 333

Spr

ing

Tub

e C

onst

ruct

ion

1()

ft

%

2()

ft

%

1e%

2e%

Bas

e sh

ape:

•Lin

e•C

ircle

seg

men

t

Mirr

or b

ase

shap

e 2x

“Wra

p”ba

se

shap

e pa

ttern

ar

ound

hel

ix

Sta

ck h

elix

pa

ttern

to

crea

te tu

be

Hal

f “n

eedl

e”

NASA/CP—2005-213655/VOL1 334

Sym

met

ry/N

ode

Cou

plin

g

All

coup

led

sets

ar

e pa

ralle

l to

the

axia

l dire

ctio

n

NASA/CP—2005-213655/VOL1 335

Nee

dle

Inte

ract

ion

Mod

els

Lin

ked

:N

eedl

es

are

“hin

ged”

with

ad

jace

nt n

eedl

es

Fre

e:A

djac

ent

need

les

have

no

influ

ence

on

allo

wed

mot

ion

Link

ed n

eedl

es a

ssum

e m

ore

need

le in

tera

ctio

n an

d in

crea

se th

e st

iffne

ss o

f the

mod

el.

NASA/CP—2005-213655/VOL1 336

Spr

ing

Tub

e A

nim

atio

n

NASA/CP—2005-213655/VOL1 337

Sin

gle

Str

and

Res

ults

Sin

gle

stra

nd6

cour

ses

per

inch

10 n

eedl

es p

er tu

rn

Sin

gle

Str

and

7 co

urse

s pe

r in

ch16

nee

dles

per

turn

Bot

h si

ngle

str

and

case

s ar

e be

st a

ppro

xim

ated

with

free

(un

linke

d) n

eedl

es

NASA/CP—2005-213655/VOL1 338

Trip

le S

tran

d R

esul

ts

•E

xper

imen

tal d

ata

show

s 2

regi

ons

–S

mal

l com

pres

sion

–le

ss in

tera

ctio

n

–H

igh

com

pres

sion

–m

ore

inte

ract

ion

•Li

ne s

lope

s–

Link

ed m

odel

: 0.0

07 lb

f/in.

–H

igh

com

pres

sion

0.0

07 lb

f/in.

–U

nlin

ked

mod

el: 0

.004

lbf/i

n.

–Lo

w c

ompr

essi

on 0

.005

lbf/i

n.

Hig

her

wir

e d

ensi

ty s

pri

ng

tu

bes

ex

hib

it m

ore

nee

dle

inte

ract

ion

Trip

le s

tran

d4.

9 co

urse

s pe

r in

ch10

nee

dles

per

turn

NASA/CP—2005-213655/VOL1 339

Sum

mar

y of

Pre

load

Dev

ice

Ana

lyse

s

•B

oth

mod

els

verif

ied

with

exp

erim

enta

l dat

a

•C

ante

d C

oil S

prin

g–

Ene

rgy

stor

age

is in

ben

ding

–In

clud

ing

a fr

ictio

n m

odel

and

com

parin

g w

ith a

long

er te

st

sprin

g im

prov

es a

ccur

acy

•S

prin

g tu

be–

Line

ar b

ehav

ior

pred

icte

d

–“F

ree”

need

les

mod

el is

bet

ter

for

less

den

se tu

bes

–C

ombi

natio

n of

“fr

ee”

and

linke

d m

odel

s yi

elds

bes

t pre

dict

ion

for

mor

e de

nse

sprin

g tu

bes

•B

oth

mod

els

guid

e an

d op

timiz

e de

sign

of h

igh

tem

pera

ture

sea

l pre

load

dev

ices

NASA/CP—2005-213655/VOL1 340

HIGH-TEMPERATURE METALLIC SEAL DEVELOPMENT FOR AERO PROPULSION AND GAS TURBINE APPLICATIONS

Greg More

Advanced Products North Haven, Connecticut

Amit Datta

Advanced Components and Materials East Greenwich, Rhode Island

High Temperature Metallic High Temperature Metallic Seal Development For Seal Development For

Aero Propulsion and Gas Aero Propulsion and Gas Turbine ApplicationsTurbine Applications

Greg More, Advanced ProductsNorth Haven, CT

Dr. Amit Datta, Advanced Components & MaterialsEast Greenwich, RI

NASA/CP—2005-213655/VOL1 341

Ou

tlin

e■

Bac

kgro

un

d

■In

no

vati

ve s

eal d

esig

n w

ith

tu

rbin

e b

lad

e al

loy

spri

ng

an

d it

s p

erfo

rman

ce

■N

ew c

on

cep

t w

ith

sin

gle

cr

ysta

l bla

de

allo

y sp

rin

gs

■C

on

clu

sio

ns

NASA/CP—2005-213655/VOL1 342

Bac

kgro

un

d:

Str

ess

rela

xati

on

of

curr

ent

cold

fo

rmab

le s

eals

Sea

l gap

is c

reat

ed r

esu

ltin

g f

rom

str

ess

rela

xati

on

at

elev

ated

tem

per

atu

res.

Th

e o

rig

inal

sea

l hei

gh

t h

o is

red

uce

d t

o h

c cr

eati

ng

a g

ap w

hen

th

e fl

ang

e m

ove

s aw

ay f

rom

th

e co

mp

ress

ed c

on

dit

ion

.

NASA/CP—2005-213655/VOL1 343

Bac

kgro

un

d:

Inef

fici

ency

ass

oci

ated

wit

h

curr

ent

seal

s

■E

xpen

sive

ble

ed a

ir is

nee

ded

fo

r co

olin

g c

urr

ent

seal

s if

u

sed

ab

ove

130

0ºF

. Les

s co

mp

ress

ed a

ir is

ava

ilab

le t

o

gen

erat

e p

ow

er!

■S

eals

are

pro

tect

ed f

rom

hig

h t

emp

erat

ure

en

viro

nm

ents

w

ith

co

mp

lex

stru

ctu

res,

incr

easi

ng

wei

gh

t an

d c

ost

NASA/CP—2005-213655/VOL1 344

Bac

kgro

un

d:

Mar

ket

Nee

ds

for

Hig

h

Tem

per

atu

re(1

500

-18

00ºF

)Sea

ls o

r S

pri

ng

s

■R

edu

ce c

oo

ling

air

an

d s

imp

lify

seal

cav

ity

st

ruct

ure

in c

urr

ent

gas

tu

rbin

e en

gin

es.

■G

as T

urb

ine

Co

mb

ust

or

ap

plic

atio

ns

to h

old

th

e C

MC

lin

er.

■S

pri

ng

load

ing

dev

ice

for

extr

emel

y h

igh

(>20

00ºF

)te

mp

erat

ure

cer

amic

slid

ing

sea

ls.

■S

eals

fo

r en

gin

e h

eat

exch

ang

er p

anel

s an

d b

ack-

up

st

ruct

ure

s fo

r fu

ture

hyp

erso

nic

en

gin

es.

NASA/CP—2005-213655/VOL1 345

Inn

ova

tive

Sea

l wit

h B

lad

e A

lloy

Sp

rin

g

Th

in c

old

fo

rmab

le a

lloy

jack

et p

rovi

din

g a

co

nti

nu

ou

s se

alin

g s

urf

ace

Cas

t b

lad

e al

loy

spri

ng

en

erg

izer

fo

r o

per

atio

n u

p t

o 1

800

oF

MA

RM

247

Sp

rin

g

Hay

nes

214

Jac

ket

NASA/CP—2005-213655/VOL1 346

Inn

ova

tive

Sea

l wit

h B

lad

e A

lloy

Sp

rin

g

Cas

t B

lad

e al

loys

hav

e ex

trem

ely

hig

h s

tren

gth

3575

1600

Was

pal

oy,

po

ly

crys

tal

1010

014

72IN

CO

718

,po

ly

crys

tal

1811

416

00C

MS

X4,

Sin

gle

cr

ysta

l

1213

014

00M

AR

M 2

47, p

oly

cr

ysta

l

Elo

ng

atio

n,%

Yie

ld S

tren

gth

,ksi

Tem

per

atu

re,º

FA

lloy

Bla

de

allo

ys a

lso

hav

e su

per

ior

cree

p a

nd

str

ess

rup

ture

str

eng

th

com

par

ed t

o c

old

fo

rmab

le s

up

eral

loys

. H

ence

, bla

de

allo

ys h

ave

hig

her

res

ista

nce

to

str

ess

rela

xati

on

.

Bla

de

allo

ys a

re a

vaila

ble

on

ly in

th

e ca

st c

on

dit

ion

( p

oly

or

sin

gle

cr

ysta

l )

NASA/CP—2005-213655/VOL1 347

Inn

ova

tive

Sea

l wit

h B

lad

e A

lloy

Sp

rin

gC

om

par

iso

n o

f h

igh

tem

per

atu

re s

ealin

g d

esig

n

Sta

nd

ard

E-S

eal w

ith

bla

de

allo

y sp

rin

g

Cro

ss s

ecti

on

Sta

nd

ard

E-S

eal p

rod

uce

d

fro

m h

igh

tem

per

atu

re

Was

pal

oy

allo

y

Cro

ss s

ecti

on

NASA/CP—2005-213655/VOL1 348

Inn

ova

tive

Sea

l wit

h B

lad

e A

lloy

Sp

rin

g

Exp

erim

enta

l pro

ced

ure

:■

Sea

ls w

ere

com

pre

ssed

20%

bet

wee

n f

lan

ges

an

d h

eate

d t

o 1

600

oF

fo

r va

rio

us

tim

e p

erio

ds

■A

fter

eac

h e

xpo

sure

, sea

ls w

ere

coo

led

to

ro

om

te

mp

erat

ure

to

mea

sure

ch

ang

e in

sea

l fre

e h

eig

ht

■S

eals

wer

e le

ak t

este

d t

o e

valu

ate

hig

h

tem

per

atu

re e

xpo

sure

on

sea

ling

per

form

ance

■B

oth

fre

e h

eig

ht

chan

ge

and

sea

l lea

kag

e w

ere

plo

tted

as

a fu

nct

ion

of

exp

osu

re t

ime

at 1

600

oF

NASA/CP—2005-213655/VOL1 349

Inn

ova

tive

Sea

l wit

h B

lad

e A

lloy

Sp

rin

gC

om

par

iso

n o

f se

al f

ree

hei

gh

t ch

ang

e vs

. exp

osu

re t

ime

at 1

600F

NASA/CP—2005-213655/VOL1 350

Inn

ova

tive

Sea

l wit

h B

lad

e A

lloy

Sp

rin

g

Co

mp

aris

on

of

seal

leak

age

vs. e

xpo

sure

tim

e at

160

0F

NASA/CP—2005-213655/VOL1 351

New

Des

ign

wit

h s

ing

le c

ryst

al s

pri

ng

■C

ast

po

lycr

ysta

llin

e M

AR

M24

7 ri

ng

was

mac

hin

ed

usi

ng

ED

M -

--p

roh

ibit

ivel

y ex

pen

sive

!■

By

con

tro

llin

g f

ing

er d

esig

n(

ang

le, t

hic

knes

s, w

idth

, #o

f fi

ng

ers/

linea

r in

ch…

) sp

rin

g r

ate

or

seal

ing

load

co

uld

be

adju

sted

■S

ing

le c

ryst

als

are

mo

re r

esis

tan

t to

cre

ep a

nd

st

ress

rel

axat

ion

bec

ause

of

the

elim

inat

ion

of

gra

in

bo

un

dar

ies

■S

ing

le c

ryst

al f

ing

ers

can

be

cast

to

nea

r–n

et s

hap

e an

d a

ttac

hed

to

eas

ily m

ach

inab

le a

nd

ch

eap

er

sup

eral

loy

rin

g

NASA/CP—2005-213655/VOL1 352

New

Des

ign

wit

h s

ing

le c

ryst

al

spri

ng

Co

st e

ffec

tive

des

ign

with

cas

t si

ng

le c

ryst

al b

lad

e al

loy

spri

ng

s at

tach

ed t

o lo

wer

co

st m

ach

ined

rin

g

NASA/CP—2005-213655/VOL1 353

■H

igh

est

stre

ss c

om

po

nen

ts(

fin

ger

s) a

re m

ade

fro

m

the

mo

st c

reep

res

ista

nt

sin

gle

cry

stal

s, c

ost

ef

fect

ivel

y ca

st t

o n

et s

hap

e

■In

div

idu

al f

ing

ers

are

atta

ched

to

a lo

wer

co

st

sup

eral

loy

rin

g w

hic

h a

re e

asily

fab

rica

ted

. Str

esse

s in

th

e ri

ng

are

sig

nif

ican

tly

low

er

■D

evic

e sp

rin

g r

ate

can

be

easi

ly c

han

ged

by

sele

ctin

g t

he

# fi

ng

er/li

nea

r in

ch a

nd

th

e fi

ng

er

des

ign

New

Des

ign

wit

h s

ing

le c

ryst

al s

pri

ng

NASA/CP—2005-213655/VOL1 354

■B

y se

lect

ing

“so

ft”

crys

tallo

gra

ph

ic a

xis,

<10

0>,

alo

ng

th

e sp

rin

g a

xis,

th

e el

asti

c ra

ng

e o

f d

efle

ctio

n

can

be

max

imiz

ed

■F

or

Ni,

E11

1/ E

100

= 2.

2

■E

last

ic r

ang

e o

f se

al h

eig

ht

chan

ge

or

seal

def

lect

ion

ca

n b

e m

ore

th

an d

ou

ble

d b

y se

lect

ing

cry

stal

o

rien

tati

on

■F

EA

op

tim

izat

ion

has

yie

lded

des

ign

s w

ith

ela

stic

ra

ng

es m

ore

th

an 3

0% o

f se

al h

eig

ht

at 1

600-

1800

ºF

New

Des

ign

wit

h s

ing

le c

ryst

al s

pri

ng

NASA/CP—2005-213655/VOL1 355

Maximum Stress in Finger Spring

Def

lect

ion

or

Ch

ang

e in

Sea

l Hei

gh

t

Fin

ger

Axi

s II

to <

111>

Fin

ger

Axi

s II

to <

100>

1

2

Infl

uen

ce o

f C

ryst

allo

gra

ph

ic D

irec

tio

n

on

Sea

l Def

lect

ion

Cap

abili

ties

NASA/CP—2005-213655/VOL1 356

New

Des

ign

wit

h s

ing

le c

ryst

al s

pri

ng

Ap

plic

atio

ns

oth

er t

han

sea

ls

Tra

nsi

tio

n f

aste

ner

bet

wee

n m

etal

an

d c

eram

ic

com

po

nen

ts w

ith

a la

rge α-

mis

mat

ch

Co

mb

ust

or

CM

C li

ner

–lo

w lo

ad, l

arg

e d

efle

ctio

n

spri

ng

at

1800

F

Lo

w-l

oad

/ hig

h d

efle

ctio

n s

pri

ng

en

erg

izer

fo

r ex

trem

ely

hig

h t

emp

erat

ure

(>2

000F

) ce

ram

ic

slid

ing

sea

l

Cer

amic

R

ope

Sea

l

Zirc

oniu

m S

eal

Inte

rfac

e

Sin

gle

Cry

stal

B

lade

Allo

y S

prin

g

Met

al

Cas

ing

Cer

amic

Lin

er

Hig

h T

empe

ratu

re

Ann

ular

Spr

ing

NASA/CP—2005-213655/VOL1 357

Co

ncl

usi

on

s

■A

n in

no

vati

ve s

eal c

on

sist

ing

of

a b

lad

e al

loy

spri

ng

en

erg

izer

an

d c

old

fo

rmab

le s

up

eral

loy

jack

et c

an b

e d

esig

ned

fo

r ap

plic

atio

ns

abo

ve 1

500

oF

■T

he

seal

is e

xpec

ted

to

min

imiz

e co

olin

g a

ir a

nd

leak

age

ther

eby

enh

anci

ng

en

gin

e ef

fici

ency

■T

he

des

ign

can

be

furt

her

imp

rove

dw

ith

cas

t n

et s

hap

e si

ng

le c

ryst

al s

pri

ng

s at

tach

ed t

o a

n e

asily

mac

hin

able

su

per

allo

y ri

ng

■T

he

sin

gle

cry

stal

sp

rin

g s

tru

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NASA/CP—2005-213655/VOL1 358

OXIDATION OF HIGH-TEMPERATURE ALLOY WIRES FOR HYBRID SEAL APPLICATIONS

Elizabeth J. Opila


Cleveland, Ohio

Jonathan A. Lorincz CON/SPAN

Dayton, Ohio

Marissa M. Reigel Colorado School of Mines

Golden, Colorado

Jeffrey J. Demange University of Toledo

Toledo, Ohio

Small diameter wires (150 to 250 μm) of the high-temperature alloys Haynes 188, Haynes 230, Haynes 214, Kanthal A1 and PM2000 were oxidized at 1204 °C in dry oxygen or 50 percent H2O/50 percent O2 for 70 hours. The oxidation kinetics were monitored using a thermogravimetric technique. Additional cyclic oxidation exposures were conducted in air for one hour cycles at 1204 °C for times up to 70 hours. Oxide phase composition and morphology of the oxidized wires were determined by x-ray diffraction, field emission scanning electron microscopy, and energy dispersive spectroscopy. The alumina-forming alloys, Kanthal A1 and PM2000, out-performed the chromia-forming alloys under these test conditions. Correlations between oxidation lifetime and wire diameter were considered. PM2000 was recommended as the most promising candidate for advanced hybrid seal applications for space reentry control surface seals or hypersonic propulsion system seals.

Glenn Research Center at Lewis Field

Oxidation of High-Temperature Alloy Wires for Hybrid Seal Applications

Elizabeth J. Opila, Jonathan A. Lorincz*, Marissa M. Reigel^, Jeffrey J. DeMange&

NASA Glenn Research Center, Cleveland, Ohio*CON/SPAN, Dayton, OH

^Colorado School of Mines, Golden, CO&University of Toledo, Toledo, OH

2004 NASA Seal/Secondary Air System WorkshopOhio Aerospace Institute

November 10, 2004

NASA/CP—2005-213655/VOL1 359


Why study oxidation of wires?Understanding of wire oxidation needed for development of advanced high-temperature seals for future hypersonic and reentry vehicles.

•Structural seal restricts hot gas leakage to underlying low-temperature control actuators

• Wire overwrap needed for wear resistance

NASA/CP—2005-213655/VOL1 360


State of the art hybrid seals

• Tmax=800°C• Ceramic batting or fiber core (alumina or

aluminosilicate)

• Haynes 188 or 230 overwrap/overbraid, 40-125 micron diameter wire

Nextel 312 fibersHaynes 230 overbraid

NASA/CP—2005-213655/VOL1 361


Requirements for advanced hybrid seals

• Withstand temperatures up to 1400°C• Operate without active cooling• Flexible, resilient, wear resistant

• Airframe control surface seals– Reentry environment: reduced pressure air, plasma– Reusability of 10 to 100 cycles of 30 minutes each

• Hypersonic propulsion system seals– Propulsion environment: high pressure water vapor– Reusability of 1000 cycles of 250 sec/cycle (70 hours)

NASA/CP—2005-213655/VOL1 362


Objectives of this study

• Characterize isothermal and cyclic oxidation resistance of high-temperature alloy wires

• Guide selection of higher temperature alloys for hybrid seal applications

NASA/CP—2005-213655/VOL1 363


Alloy wires

Fe base, 20 Cr, 5.5 Al, 0.5 Ti, 0.5 Y2O3

Fe base, 22 Cr, 5.8 Al, Si<0.7, Mn<0.4

Ni base, 16 Cr, 4.5 Al, 3 Fe, 0.2 Si, trace: Mn, Zr, C, B, Y

Ni base, 22 Cr, 14 W, Co<5, Fe<3, 2 Mo, 0.5 Mn, 0.4 Si, 0.3 Al, trace: C, La, B

Co base, 22 Cr, 22 Ni, 14 W, Fe< 3, Mn <1.25, 0.5 Si, 0.12 La, trace: C, B

Composition, wt%

150

250

250

250

250

Diameter, μm

PM 2000

Kanthal A1

Haynes 214

Haynes 230

Haynes 188

Alloy

Haynes 188, and 230 are chromia forming alloys. Haynew 214 is a marginal alumina forming alloy. Kanthal A1 and PM2000 are alumina forming alloys. PM2000 contains the reactive element Y2O3 which improves oxide adherence.

Note smaller diameter wire for PM2000.

NASA/CP—2005-213655/VOL1 364


Experimental Procedure

• Coil 60 cm length of wire• Oxidation of wires

– 1204°C (2200°F)– 1 atm– 70h– Three exposure environments

• Isothermal: O2 (0.4 cm/sec) • Isothermal: 50% H2O/50% O2 (4.4

cm/sec)• Cyclic: stagnant air, 1h hot, 20

min. cool, 70 cycles

• Weight change of wires• X-ray diffraction (XRD) analysis• Field Emission-Scanning Electron

Microscopy/Energy Dispersive Spectroscopy (FE-SEM/EDS)

0.5 cm

Weight gain indicates oxide formation. Weight loss indicates loss of oxide by spallation or volatilization.

NASA/CP—2005-213655/VOL1 365


Thermogravimetric Analysis

NASA/CP—2005-213655/VOL1 366


Cyclic Oxidation

One cycle = 1 hour hot, 20 minute cool..

NASA/CP—2005-213655/VOL1 367


Results for wire oxidation in dry O2

NASA/CP—2005-213655/VOL1 368


Weight change of wire: 1204°C, dry O2, 0.4 cm/s

0 10 20 30 40 50 60 70

time, h

0

2

4

6

8

10

12

14

16

18

20

spec

ific

wei

ght c

hang

e, m

g/cm

2Haynes 188Haynes 230Haynes 214Kanthal A1Kanthal A1PM 2000PM 2000

Haynes 188 completely oxidized after about 20 hours.

NASA/CP—2005-213655/VOL1 369


Weight change of wire: 1204°C, dry O2, 0.4 cm/s

0 10 20 30 40 50 60 70

time, h

0.0

0.5

1.0

1.5

2.0

2.5

3.0

3.5

4.0

spec

ific

wei

ght c

hang

e, m

g/cm

2Haynes 230

Haynes 214

PM2000

Kanthal A1

Haynes 188

Note: All compositions except PM2000 exhibited spallation on cooling

Haynes 230, PM2000, Kanthal A1 show desired parabolic oxidation kinetics.

Haynes 214 shows protective oxidation after initial transient of fast oxidation.

Bump in PM2000 kinetics at 45h may be due to depletion of alumina and beginning of chromia scale formation.

NASA/CP—2005-213655/VOL1 370


Coiled wires after oxidation: 1204°C, 70h, dry O2, 0.4 cm/s

Haynes 188 Haynes 230 Haynes 214

PM 2000Kanthal A1

0.5 cm

Haynes 188 brittle. Haynes 230, 214, Kanthal A1 show spallation to bare metal.

NASA/CP—2005-213655/VOL1 371


X-ray diffraction results for wires:1204°C, 70h, dry O2, 0.4 cm/s

Fe base, 20 Cr, 5.5 Al, 0.5 Y2O3

Fe base, 22 Cr, 5.8 Al

Ni base, 16 Cr, 4.5 Al

Ni base, 22 Cr, 14 W

Co base, 22 Cr, 22 Ni, 14 W

Major components of alloy, wt%

α Al2O3

α Al2O3

NiCr2O4/NiAl2O4NiO

Cr2O3NiCr2O4

NiO

NiWO4NiCr2O4

NiO

Oxidation products

PM 2000

Kanthal A1

Haynes 214

Haynes 230

Haynes 188

Alloy

Chromia formation not observed for Haynes 188 after 70h. Note presence of NiWO4 oxide phase.

First three alloys show spinel (AB2O4) formation. This is not a protective oxide.

NASA/CP—2005-213655/VOL1 372


Cross-sectional FE-SEM of wire: 1204°C, 70h, dry O2, 0.4 cm/s

100 μm

Haynes 214

100 μm

Haynes 230

100 μm

Kanthal A1

100 μm

Haynes 188

100 μm

PM2000

Haynes 188: (Co,Ni)O, CoCr2O4, CoWO4-bright phase. Wire completely consumed.

Haynes 230: Cr2O3

Haynes 214: Al2O3 inner layer, spinel outer layer, NiO blocks on surface

Kanthal A1: Oxide scale completely nonadherent.

NASA/CP—2005-213655/VOL1 373


Summary of results:Wire oxidation 1204°C, 70h, dry O2, 0.4 cm/s

• Haynes 188 completely oxidized• Haynes 230 formed chromia scale, spalled on cool down• Haynes 214 formed inner alumina scale, external spinel, NiO,

spalled on cool down• Kanthal A1 formed alumina scale, completely nonadherent after

cool down• PM2000 smaller diameter wire formed an adherent alumina

scale, aluminum completely depleted from wire, inner discontinuous layer of chromia, internal void formation

NASA/CP—2005-213655/VOL1 374


Results for wire oxidation in 50% H2O/50% O2

NASA/CP—2005-213655/VOL1 375


0 10 20 30 40 50 60 70

time, h

0

1

2

3

4

5

6

7

8

spec

ific

wei

ght c

hang

e, m

g/cm

2Haynes 188Kanthal A1PM2000

Weight change of wire: 1204°C, 50% H2O/50 %O2, 4.4 cm/s

Different weight change behavior for Haynes 188 explained on next slide.

NASA/CP—2005-213655/VOL1 376


0 10 20 30 40 50 60 70

time, h

0

2

4

6

8

10

12

14

16

18

20

spec

ific

wei

ght c

hang

e, m

g/cm

2

dry O250% O2/50% H2O

wire completely oxidized

oxid

atio

n

oxid

atio

n &

vola

tiliz

atio

n

volatilization: WO2(OH)2(g), Ni(OH)2(g)

Weight change of Haynes 188 wire: 1204°C

NASA/CP—2005-213655/VOL1 377


Volatility of oxides formed on Haynes 188 exposed at 1204°C, 50% H2O/50 %O2

Deposit found on hanger downstream of sample. Deposit determined to be WO3and NiWO4 by XRD.

Calculated gas species formed from oxides in water vapor. W is very volatile.

NASA/CP—2005-213655/VOL1 378


0 10 20 30 40 50 60 70

time, h

0.0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

1.6

1.8

2.0

spec

ific

wei

ght c

hang

e, m

g/cm

2

Kanthal A1PM2000

Weight change of wire: 1204°C, 50% H2O/50 %O2, 4.4 cm/s

Breakaway oxidation for PM2000 at 68 h.

NASA/CP—2005-213655/VOL1 379


0 10 20 30 40 50 60 70

time, h

0.0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

1.6

1.8

2.0

spec

ific

wei

ght c

hang

e, m

g/cm

2

dry O2dry O250% O2/50% H2O

Weight change of PM2000 wire: 1204°C

NASA/CP—2005-213655/VOL1 380


Correlation between breakaway oxidation and wire diameter

• Breakaway oxidation occurs when aluminum is depleted to low level: protective oxide scale formation no longer possible

• Aluminum depletion depends on ratio of wire volume to surface area.– Extrapolating time to breakaway for PM2000 at 1204°C in 50% H2O for

other wire diameters.

Wm = weight of metal consumed during oxidation, V = wire volume, A = wire surface area, r = wire radius, t = oxidation time

– Experimentally determined time to breakaway for 150 μm dia wire is 68h.

– Predict time to breakaway for 250 μm dia wire is 189h for comparison to other alloy wires of 250 μm dia.

– Predict time to breakaway for 40 μm dia wire at 5h for finest diameter wire proposed for use in seal applications.

2b

2a

b

a

mAV

m

r

rtt

tWrW

=

∝=∝

NASA/CP—2005-213655/VOL1 381


Fe base, 20 Cr, 5.5 Al, 0.5 Y2O3




(Fe,Cr)2O3

α Al2O3

NiO NiCr2O4

Oxidation products

PM 2000

Kanthal A1

Haynes 188

Alloy

X-ray diffraction results for wires:1204°C, 70h, 50% H2O/50% O2, 4.4 cm/s

No W containing phase found on Haynes 188. W is completely volatilized.

Al2O3 not found on PM2000 after 70h in this environment in contrast to results in dry O2.

NASA/CP—2005-213655/VOL1 382


Cross-sectional FE-SEM of wire: 1204°C, 70h, 50% H2O/50% O2, 4.4 cm/s

100 μm

100 μm

100 μm

Haynes 188

Kanthal A1

PM2000

Porosity in Haynes 188 where NiWO4 found after exposure in dry O2.

Oxide scale not adherent on Kanthal A1.

Adjacent cross-sections of PM2000 show one completely oxidized, the other metallic with nearly protective scale still intact.

NASA/CP—2005-213655/VOL1 383


Summary of results: Wire oxidation, 1204°C, 70h, 50% H2O/50% O2, 4.4 cm/s

• Haynes 188 completely oxidized, all W volatilized leaving porosity.

• Kanthal A1 formed alumina scale, completely nonadherent, much of scale spalled.

• PM2000 smaller diameter wire experienced breakaway oxidation. Portions of wire completely oxidized, show remainderof alumina scale. Portions of wire show protective alumina scale.

• Time to breakaway oxidation expected to vary with the square of the wire diameter.

NASA/CP—2005-213655/VOL1 384


Cyclic Oxidation Results

NASA/CP—2005-213655/VOL1 385


Weight change of wire: 1204°C, stagnant air, 1 h cycles

0 10 20 30 40 50 60 70

Time (h)

0

2

4

6

8

10

Spe

cific

wei

ght c

hang

e (m

g/cm

2 )

Haynes 188Haynes 188PM2000PM2000Kanthal A1Kanthal A1failure

Haynes 188 weight change similar to isothermal exposures in that oxidation is about complete at 20h.

PM2000 showing parabolic oxidation to 60+ hours.

Kanthal A1 shows weight loss and spallation after third cycle. Breakaway oxidation at 55h.

NASA/CP—2005-213655/VOL1 386


X-ray diffraction results for wires:1204°C, 1h cycles, stagnant air

>70

60

30

Time to failure, h

Fe base, 20 Cr, 5.5 Al, 0.5 Y2O3




tbd

Fe2O3

(Ni,Co)O (Ni,Co)Cr2O4

Oxidation products

PM 2000

Kanthal A1

Haynes 188

Alloy

Non protective Fe2O3 found for Kanthal A1.

NASA/CP—2005-213655/VOL1 387


Cross-sectional FE-SEM of wire: 1204°C, 1h cycles, stagnant air

100 μm

100 μm

100 μm

Haynes 188, 30 cycles

Kanthal A1, 60 cycles

Haynes 188 shows core of wire still present after 30 cycles.

Kanthal A1 has some cross-sections just beginning non-protective oxide formation (top) and others completely oxidized (bottom). The original wire diameter of Kanthal A1 is marked by darker phases (alumina and chromia) in lower right micrograph.

NASA/CP—2005-213655/VOL1 388


• Haynes 188 had similar failure time as in isothermal testing.

• Kanthal A1 failed earlier in cyclic oxidation testing. Weight loss and oxide spallation detected as early as third temperature cycle. Breakaway oxidation occurred after 55 cycles.

• PM2000 showed parabolic oxidation kinetics throughout 70h cyclic test. No evidence of spallation.

Summary of results: Wire oxidation, 1204°C, 1h cycles, stagnant air

NASA/CP—2005-213655/VOL1 389


Conclusions

• Alumina-forming alloys with reactive element additions perform best at 1204°C under all test conditions: O2, H2O, temperature cycling– Slow growing oxide– Alumina is the most stable protective oxide scale in water vapor

– Adherent scales

• Small diameter wires have limited oxidation lifetimes. – Limited reservoir of aluminum for protective scale formation

– Smaller diameter wires more prone to spallation: increased stress in oxide due to larger radius of curvature

• Oxidation lifetimes can be predicted based on the wire diameter and rate of aluminum loss.

• From these results PM2000 is recommended as the best candidate for further development in advanced hybrid seal applications.

NASA/CP—2005-213655/VOL1 390


Acknowledgments

Jim Smialek, many helpful discussionsRalph Garlick, X-ray diffraction

Patrick Dunlap, Bruce Steinetz, program support

This work was funded by the Next Generation Launch Technology Program

NASA/CP—2005-213655/VOL1 391

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NASA/CP—2005-213655/VOL1 394

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National Aeronautics and Space AdministrationJohn H. Glenn Research Center at Lewis FieldCleveland, Ohio 44135–3191

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October 2005

NASA CP—2005-213655–VOL1

E–15144–1

WBS–22–714–70–42

400

2004 NASA Seal/Secondary Air System Workshop

Bruce M. Steinetz and Robert C. Hendricks, editors

Seals; Turbine; Clearance control, Materials; Analyses; Experimental; Design, Dockingmechanism; Leakage

Unclassified -UnlimitedSubject Categories: 37, 16, and 99 Distribution: Nonstandard

Proceedings of a conference held at Ohio Aerospace Institute sponsored by NASA Glenn Research Center, Cleveland,Ohio, November 9–10, 2004. Responsible person, Bruce M. Steinetz, organization code RSM, 216–433–3302.

The 2004 NASA Seal/Secondary Air System workshop covered the following topics: (i) Overview of NASA’s new ExplorationInitiative program aimed at exploring the Moon, Mars, and beyond; (ii) Overview of the NASA-sponsored Ultra-Efficient EngineTechnology (UEET) program; (iii) Overview of NASA Glenn’s seal program aimed at developing advanced seals for NASA’sturbomachinery, space, and reentry vehicle needs; (iv) Reviews of NASA prime contractor and university advanced sealing conceptsincluding tip clearance control, test results, experimental facilities, and numerical predictions; and (v) Reviews of material developmentprograms relevant to advanced seals development. The NASA UEET overview illustrated for the reader the importance of advancedtechnologies, including seals, in meeting future turbine engine system efficiency and emission goals. For example, the NASA UEETprogram goals include an 8- to 15-percent reduction in fuel burn, a 15-percent reduction in CO2, a 70-percent reduction in NOx, CO,and unburned hydrocarbons, and a 30-dB noise reduction relative to program baselines. The workshop also covered several programsNASA is funding to develop technologies for the Exploration Initiative and advanced reusable space vehicle technologies. NASA planson developing an advanced docking and berthing system that would permit any vehicle to dock to any on-orbit station or vehicle, aspart of NASA’s new Exploration Initiative. Plans to develop the necessary mechanism and “androgynous” seal technologies werereviewed. Seal challenges posed by reusable re-entry space vehicles include high-temperature operation, resiliency at temperature toaccommodate gap changes during operation, and durability to meet mission requirements.

2004 NASA Seal/Secondary Air System Workshop

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