Tugas Besar Kelompok 13 - Supersonic Intake Design

7/25/2019 Tugas Besar Kelompok 13 - Supersonic Intake Design

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FINAL TASK REPORT

SUPER SONIC INTAKE DESIGN

AE3210 AERODYNAMICS 2

By:

GROUP 13

Bintang A.S.W.A.M [13613019]

Irvan Hakim F. [13613025]

Muhammad Fazlur R [13613039]

M. Rafiqi Sitompul [13613057]

Due Date : Wednesday, 18th May 2016

AERONAUTICS AND ASTRONAUTICS MAJOR

FACULTY OF MECHANICAL AND AEROSPACE

ENGINEERING

INSTITUTE TECHNOLOGY OF BANDUNG

2016



Supersonic Intake Design

Department of Aerospace Engineering, Bandung Institute of Technology,

Indonesia

Abstract: This research’s goal is to perform aerodynamic design of an intake for

supersonic fighter aircraft. The intake is design to be have the most efficient

external-compression inlet. The analysis is taken at one flight condition that is

operate at 41000 ft. The air is assumed to be inviscid and calorically perfect. The

flow also assumed to be two dimensional and for the nominal design we will assume

the plane is flying level (angle of attack = 0). The calculations will be done

assuming gamma = 1.4. The Mach number after the normal shock at the entrance

to the subsonic diffuser is required to be M = 0.8 and the maximum angle that the

flow can turn prior to entering the diffuser is 25 degrees. The range of Mach

numbers to be considered from 3.1 ≤ ≤ 3.5.

Key words:

External-compression, inviscid, efficient

Introduction

Aircraft which is flying at supersonic speed or above the speed of sound has

a lot of problems for engineers. One of the problems faced is the design of Intake.

The Intake is a duct before the inlet of compressor. Because of the combustion

process occurs quite well at subsonic velocity air flow. Therefore, the incoming air

flow in Combustion Engines are required to become subsonic. To meet these

requirements, the flow must be slowed from supersonic to less than Mach 1.

In this project, the author assumes that ideally the flow enters the divergent

nozzle at the front of the engine intake should be about M = 0,8. The high number

Mach is reduced in divergent nozzle for the special needs of the machine. In this

project the author will design some design Intake for Mach numbers between M =

3,1 and M = 3,5. The final design is expected to be able to slow down the flow into

M = 0.8 and also have the highest Pressure Recovery.



The aircraft engine requires a supply of uniform hogt total pressure recovery

air for good performance. It also related to total pressure loss which effects to the

engine thrust and consequently the fuel consumption. Thefore it is important to

maximaze the total pressure recovery at the inlet of engine. The total pressure

recovery is noted as ratio of the total pressure of the airflow before entering the

compressor to the freestream.

Proposed Method and Implementation

The preliminary design is divided into two following subpart. The first task is to

determine the number of ramps or oblique shocks and then select the subsonic

diffuser. The second task is to determine the total pressure recovery of the selected

design and choose the most efficient design.

A. Inlet Configuration and Properties

1. Inlet Preliminary Design

In preliminary design, we try to deciding the basic configuration of the inlet

such as theta angle. After that we calculated the Oblique Shock Wave (OSW)

properties for each stage interrelated so that in the end of the stage we have a normal

shock wave. The flight atmosphere condition at 41000 feet from ISA has static

pressure 17864.42 Pa, Temperature 216.65 K.

Assumption in calculation :

Adiabatic

No slip wall

No relative pressure along farfield and inlet

a. Using Manual Iteration Method

The preliminary design first and second Inlet (Inlet 1 & 2) is informed in

this table below.



Table 1 inlet preliminary design 1

inlet

first

OSW

Secont

OSW

Third

OSW

Fourth

OSW

fifth

OSW

Normal

Shock

theta 10 10 10 10 8

beta 25.46 29.4 34.3885 41.31 49.9651

mach number 3.3 2.7468 2.2995 1.9077 1.5537 1.2692 0.8


Beta angle is calculated by doing iteration for known theta. The relation between

theta, beta and mach number for OSW is noted below

θ = flow deflection angle

β = oblique shock wave angle

γ = the ratio of specific heat

M1 = Mach number in fort of oblique shockwave

The geometry of oblique shock wave is demonstrated as the figure below

Figure 1. oblique shock wave geometry



We have to know the Mach number behind the oblique shock wave (M2) for the

next stage oblique calculation. M2 can be determined by the formula :

Pressure static ratio between behind and in front of oblique is calculated by

To calculate the total pressure ratio, we need to know the ratio of temperature. With

isentropic assumption we can determine the total pressure. To get the temperature ratio

we have to calculate the density property of air flow.

Density ratio :

Total pressure ratio :

After all the properties known, Inlet Pressure Recovery (IPR),

Pi2 = Total pressure after the normal shock wave

Ptoo = Total pressure of free streeam

854919 1022180 0.8364



b. Using Optimation Approaching Method

Geometry Approach for Failed Result

Consider for Normal Shocks:

Previous Stage: M j-1, P j-1,Pt j-1, T j-1

P j-1 = Pq; Pt j-1 = Ptq; T j-1 = Tq;

Notice that Mq = M j-1 represents upstream Mach Number , meanwhile M j

represents downstream Mach Number.

Governing Equation :

1− 22− 1 ;

Next Stage: M j, P j,Pt j, T j

Consider for Oblique Shocks:

Previous Stage: Mq-1, Pq-1, Ptq-1, Tq-1 ;

Governing Equation(s) :

+ − −(+)

[−−][−+]

tan

co− ++− ;ℎ 1,2,3, … … … , 1,

Optimum criteria for flow trajectory:

− sin() ≡ ;ℎ 1,2,3, … … … … . . . , 1,

For more convenience , let define

≡ 1−sin

1; ℎ 1,2,3, … … … … . . , 1,

2 + 1 1

+ 2 ;



4 + 1

+ 1 ;

Thus,

− ;

By doing algebraic manipulation in order to obtain the results simultaneously and

also a little bit creativity to modify the governing equation, and then we’ll get

A = 19.66291154; B = 26.67264491; C= 14.32838557; (exists for

range freestream Mach Number between 1.7-5.4 )

Notice that n expresses the number of oblique shock(s) that occur.

Next Stage: Mq, Pq, Ptq, Tq ;

Equation of Total Pressure Ratio for oblique shock:

For k = 1-3,

[ +−+]

[ +−−]

Equation of Total Pressure Ratio for normal shock:

For j = k+1,

[ +−+]

[ +−−]

Thus, total pressure recovery (TPR) is

∏ × ∏

=+

=




B.

Design of The Inlet

1. 2D Geometry Modelling

The geometry of inlet is sketched using CATIA based on preliminary design we

have known.We sketch 2D inlet after that extrute becoming 3D.

Draw all theta and beta angle using CATIA

Figure design 1



Figure design 2

Obtained geometry length of the lines as follows

Figure design 1



Figure design 2

2. 3D Modelling

Extrude 2D sketch

Figure design 1



Create the Farfield, Inlet and Outlet

Figure design 1

Figure design 2

3. ANSYS Simulation



Import geometry that has been created using CATIA to ANSYS in igs. Format

Figure design 1

Create and define sections/parts such as inlet, outlet, farfield and wall.

Figure design 1



Figure design 2

Repair geometry

Figure design 1



Figure design 2

Meshing Steps

Figure design 1



Figure design 2

Figure design 1

Figure design 2



Input ti CFX, then input Boundary Condition, Domain and Initial Condition.

Figure design 1

Input condition in Inlet, Outlet, Wall and Farfield

Figure design 1



Figure design 1

Figure input condition design 2




4. Running the Simulation

Figure running design 1



C. The Inlet Design Result

Figure result design 1






Total pressure at the end = 854919 pa

Mach number = 0.8

Comparison betweet analytical result and cfd total pressure = 1:1

-Mach 3.1 free slip wall (off design)



-Mach 3.5 free slip wall(off design)

The result :

TP awal 1362560

M2 1.127735 sebelumnormal

Tp2 900350.6

mach akhir 0.8966

takhir/t2 1.078

Tp akhir 898730

p0akhir/p02 0.9982

IPR 0.659589

Table : result design 2




Failed Result

Figure Shock Detach

Figure Shock Detach



Analysis

To design an external air intake, as possible as Oblique Shock Wave hit

the cowl in order OSW not reflected in the internal intake. In the design, first sought

smallest theta and successively enlarged so that the total pressure loss can be

minimized. It happens because the total pressure before the first oblique so is high

so that the pressure loss is also high. In designing intake, as much as possible using

small angle because in addition to avoid the total pressure loss is high it can also

avoid detached-shock.

Before passing through the normal shock wave, we suggest not to

encourage area because it can thwart occurrence the normal shock wave. This case

happens if the rearing area before normal shock will accelerate the speed of the

flow.

To minimize the total pressure loss in normal, as much as possible the

flow rate approaching one. Because of the greater Mach number then normal shock

that occurs is getting stronger. To minimize the total pressure loss in the normal

shock wave, as much as possible about Mach Upstream Downstream Mach 1.1-1.3

order of about 0.81 - 0.94

Boundary layer effects caused by the walls of the intake was not overly

affect the angle beta and theta. This is evident from the results of CFD analysis has

been done, the result is the same as the analytical results that we did. As for the

boundary layer effects on speed, overall do not affect if the geometry is not too

large. However, the boundary layer remains influential on the speed around the

wall. To see a normal shock wave, is better viewed with the pressure difference.

Because if you use the speed difference, there will be the effect of the boundary

layer on the wall. So, there is the velocity distribution in the direction perpendicular

wall.



Conclusion

The design of Inlet 1 has great Inlet Pressure Recovery which is 0.8364 as

same as design 2. But, the design 1 is diffrent kind with design 2. Design 1 is two-

dimentional sigle duct, in the other hand Design 2 is two-dimentional bifurcated

duct.

Actually we have 4 design as we can see in the result of this report. But, 2

kind design has been failed so far, the problem is, first theta is to high so that it

occure shock detach. And the last design is fail because of technical problem using

ANSYS, we guess that for the Axisymmetric outward-turning has different

threarment in the meshing and define the boundary condition.

References

Anderson, John D. 1990. Modern Compressible Flow. Second Edition.

Singapore : McGraw-Hill Publishing Company

Ran, Hongjun and Dimitri Mavris. 2015. Paper : Preliminary Design of a

2D Supersonic Inlet to Maximize Total Pressure Recovery. Atlanta : Georgia

Institute of Technology

Tugas Besar Kelompok 13 - Supersonic Intake Design

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