Aerothermodynamic Study of UHTC based TPS

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Aerospace Science and Technology 9 (2005) 151–160

www.elsevier.com/locate/aesc

Aerothermodynamic study of UHTC-basedthermal protection systems

Studio aerotermodinamico di sistemi di protezionetermica basati su UHTC

Raffaele Savinoa,∗, Mario De Stefano Fumoa, Diego Paternaa, Michelangelo Serpicob

a Department of Space Science and Engineering “L.G. Napolitano”, University of Naples “Federico II”, P. le Tecchio 80, 80125 Naples, Italyb Italian Aerospace Research Center (CIRA), Via Maiorise, 81043 Capua, Italy

Received 19 August 2004; received in revised form 16 December 2004; accepted 20 December 2004

Available online 27 January 2005

Abstract

Computational Fluid Dynamics (CFD) simulations are coupled to a thermal analysis model to investigate the thermal responseUltra High Temperature Ceramics (UHTC) being considered for Thermal Protection Systems (TPS) of future reusable re-entry veh

The numerical methodology has been applied to the Sub-orbital Re-entry Test (SRT), mission foreseen in the frame of the Italianspace program. The numerical prediction of the aerodynamic heating of the “sharp” nose of the vehicle has been performed cosolution of the external aerodynamic flow field with the thermal field in the TPS; different methodologies have been applied, consideffects of surface catalysis and laminar-turbulent transition.

The results indicate that, for a correct prediction of the aerothermal loads on a UHTC material, coupled time-dependent simulboth external aerodynamic flow and internal thermal field are essential. The numerical investigations also show that surface catnegligible importance for a sharp configuration in free flight, but it must be taken into account while performing ground tests in highfacilities. Furthermore, for the considered “sharp” configuration, laminar-turbulence transition effects are critical for TPS design, smay increase TPS surface temperature of, at least, 400 K. 2005 Elsevier SAS. All rights reserved.

Resumi

In questo lavoro è stata affrontata la problematica del riscaldamento aerotermodinamico di materiali ceramici resistenti ad eleperature (UHTC Ultra High Temperature Ceramics), tra i cui possibili impieghi si prevede quello di protezione termica dei futuri verientro spaziale.

Sono state sviluppate alcune metodologie per accoppiare la soluzione del campo fluidodinamico con il campo termico nel mprotezione, considerando anche gli effetti della cataliticità superficiale e della transizione da regime laminare a turbolento.

L’analisi è stata effettuata con riferimento alla missione di rientro suborbitale (SRT) prevista nel programma spaziale italiano pnon abitati.

I risultati indicano che, per una corretta previsione dei carichi termici sui materiali UHTC, è essenziale una simulazione instazioaccoppi, istante per istante, il campo aerodinamico e quello termico nel materiale.

L’analisi degli effetti di cataliticità superficiale dimostra che essi risultano trascurabili, per configurazioni aerodinamiche “a spigonelle condizioni di volo; d’altra parte, la capacità della superficie di promuovere la ricombinazione delle specie dissociate deve essin conto durante prove sperimentali in gallerie ad alta entalpia.

* Corresponding author. Phone: +39 81 7682357, fax: +39 81 593 2044.
E-mail address:[email protected] (R. Savino).
1270-9638/$ – see front matter 2005 Elsevier SAS. All rights reserved.doi:10.1016/j.ast.2004.12.003

152 R. Savino et al. / Aerospace Science and Technology 9 (2005) 151–160

a di pro-
Inoltre, per la configurazione considerata, la transizione al regime turbolento ha un effetto critico sulla progettazione del sistemtezione termica, poiché può causare un aumento della temperatura superficiale di almeno 400 K. 2005 Elsevier SAS. All rights reserved.
Keywords:Sharp; UHTC; Radiative equilibrium; Catalysis; Transition

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1. Introduction

Much effort is currently dedicated in the aerospace comunity to study future vehicles reentering the earthmosphere. The world space agencies and industries aring to develop fully reusable space transportation systable to perform multiple launches per day in a safe, ceffective manner, without long and expensive refurbishmoperations (e.g., 40000 man-hours for Space Shuttle [1The future space transportation should resemble, fromoperational point of view, our current airports. The reusaity characteristics will insure a dramatic costs reductionmore than one order of magnitude [30]), and a simplercess to space. This aspect is of particular relevance, afterecent tragedy of Columbia and the subsequent announNASA to retire the Space Shuttle in 2010. In particular,findings of the Columbia Accident Investigation Board [1and the recent Space Exploration Initiative [9] suggestsafety be put at the top of the priorities in the design of retry vehicles.

Among possible design aspects aimed at substantlowering the reentry risks, most important is the develment of winged vehicles characterized by sharp configtions, aviation-like, with high Lift-to-Drag ratio, able to flfor long time at high altitudes and small angle of attack [1Given the entry velocity, increasing the trajectory timea much more gentle deceleration (with respect to, forample, the Space Shuttle trajectory) at high altitudes, woimprove safety because of the much lower values of theergy transferred per unit time, with much lower aerodynaheating rates and with much lower mechanical loads onstructure. This opens a completely new technical and optional scenario.

Low-risk trajectories and sharp configurations (i.e. leing edges with high curvatures) call for TPS able to wistand high temperatures. Indeed, convective heat transfcreases with the curvature, so that it is of great importato accurately predict how this heat transfer is distributedthe TPS material. At the typical orbital reentry conditioconvective heat transfer to the surface is partly conductethe solid and partly re-radiated in the atmosphere. Whesteady state is achieved, global radiative equilibrium istablished, in the sense that the (surface) overall conveheat flux is perfectly balanced by the overall surface rative flux. When conduction in the solid may be neglec(i.e. TPS materials with very low thermal conductivity), rdiative heat transfer locally balances convective heat tran(local radiative equilibrium). In both cases, the correspo
ing surface temperature is referred to as “radiative equilib-
-

f

-

rium temperature”. This temperature, depending on flconditions, may be of the order of up to 3000 K, a teperature that only very specialized materials can withstaA class of such materials is Ultra High Temperature Cerics (UHTCs) [6–8,11,15,20,23,29].

The characteristics of UHTCs, such as tensile strenghigh temperatures and oxidation resistance, allow them tused in extreme environments like those associated withpersonic flight, atmospheric reentry and rocket propulsSome compounds belonging to the UHTC family are borisuch as Zirconium Diboride (ZrB2) and Hafnium Diboride(HfB2), which have melting temperatures above 3300These materials also retain their strength (up to 150 Mat temperatures above 1500 K, exhibit good thermal retance, and can be modified with additives such as SiliCarbide (SiC) to promote oxidation resistance and improstrength.

Recent works have been dedicated to hypersonic visflows over sharp geometries and on possible improvemrelated to the utilization of UHTC materials [3,12,19,28,3

In particular, the attention has been focused on the mrials characterizations [12,19,28] or on problems relatecontinuum-rarefied transition on sharp leading edges, uboth Navier–Stokes formulations and Direct Monte Casimulations [3,31].

The present paper deals with the behaviour of sharpfigurations in the continuum regime typical of sub-orbireentry trajectories used for flight testing.

One of the objectives of this paper is to analyze differfluid-dynamic problems related to the development ofvanced TPS architectures, in particular the transient aernamic heating, and to show different methodologies to sthese problems. In particular, two different techniquesproposed, to study, with an integrated fluid-dynamic/therapproach, the time-dependent aerodynamic heating othermal protection system during the atmospheric re-ephase.

Attention is focused on a sphere-cone geometry, wsmall nose radius of curvature (Fig. 1).

Self-similar solutions of the compressible boundary lers around a stagnation point show that the stagnation pheat flux is directly dependent on flow specific total enthaand stagnation pressure, in addition to the dependencthe radius of curvature [1]. Furthermore, in hypersonic fliconditions, the vehicle kinetic energy is converted into thmal energy and, in part, into chemical energy, by the bshock wave before the body. When the shock stand offtance (Fig. 1) is relatively small, as is the case when the n
radius of curvature is small, the fraction of energy stored into

R. Savino et al. / Aerospace Science and Technology 9 (2005) 151–160 153

Nomenclature

h Heat transfer coefficient . . . . . . . . . W m−2K−1

H Total enthalpy. . . . . . . . . . . . . . . . . . . . . . . . J kg−1

M Mach numberO Mass fraction of atomic oxygenN Mass fraction of atomic Nitrogenp Pressure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pap02 Stagnation pressure behind normal shock . . . Paq Convective heat flux . . . . . . . . . . . . . . . . . W m−2

Rn Nose radius . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . mT Temperature . . . . . . . . . . . . . . . . . . . . . . . . . . . . . KT0 Total temperature . . . . . . . . . . . . . . . . . . . . . . . . . KU Surface tangential velocity component . . m s−1

Z Altitude . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . km

Greek symbols

ε Emissivityλ Thermal conductivity . . . . . . . . . . . . W m−1 K−1

ρ Density . . . . . . . . . . . . . . . . . . . . . . . . . . . . . kg m−3

σ Stefan–Boltzmannconstant . . . . . . . . . . . . . 5.67· 10−8 W m−1 K−4

Subscripts

0 Stagnation point valuese Boundary layer edgew Wall surfaces Solid

ible.clu-sis

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todes,low

, re-es,ear-toaf-

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e ofper-en-e-

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the gas chemical degrees of freedom is usually negligTherefore, in this case, the surface heat flux is almost exsively determined by Fourier diffusion, and surface catalycannot play any role. A different behaviour is expected wTPS models are tested in ground high-enthalpy facilities,to the relatively large amount of atomic species in the plaflow entering the wind tunnel test chamber. In this study,importance of chemical nonequilibrium and surface catsis effects is evaluated with regard to both free flight awind tunnel conditions.

Laminar-turbulent transition effects are well knownincrease the surface heat transfer rate. At high altitutransition and turbulence are not an issue, due to theReynolds numbers. However, at intermediate altitudesceptivity of the boundary layer to free stream disturbancsurface roughness and body vibrations allows the appance and growth of instabilities that, ultimately, leadtransition to turbulence. A number of other parametersfect transition in hypersonic flows. The transition Reyno

Fig. 1. Schematic of supersonic flow around a sphere-cone configuration.

number may be defined as a function of geometry, anglattack, Mach number, surface temperature-adiabatic temature ratio, chemistry effects and a number of minor depdencies [1]. Prediction of the transition for generic geomtries is still an open problem and a number of engineetools exist, ranging from correlation formulas to morephisticated methods, based on linear stability theoryeN

method [5]), to Direct Numerical Simulation (DNS) metods, whose application is today limited to low Reynonumbers and relatively simple geometries. For the conuration and the flight conditions investigated in the prespaper, the transition effects on heat transfer to the TPStaken into account with a semi-empirical correlation of dfor sharp cones in hypersonic conditions.

2. Reference configuration

The Unmanned Space Vehicle (USV) being developedCIRA (Italian Aerospace Research Center) [25] is an intepart of the technology program currently being pursuedItalian Aerospace Research Program (PRORA).

A schematic representation of the USV technology deonstrator is shown in Fig. 2. The primary goal of the USprogram [14,26] is to develop key technologies needesuccessfully build and operate reusable space transption systems that will significantly lower the cost of acce

Fig. 2. Vehicle geometry, coordinate system.


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Table 1USV geometric parameters

Full scale geometric parameters

Total length 8000 (m)Wing span 3.560 (m)Fuselage plan area 6.700 (m2)Wing plan area 3.744 (m2)Total area 10.444 (m2)Nose radius 0.010 (m)Wing leading edge radius 0.010 (m)Dihedral angle 5 (degBase area 0.717 (m2)

Full scale reference geometric parameters

Reference length (mean aerodynamic chord) 1.05 (mReference area 3.60 (m2)Aspect ratio 3.39 (–)

(a)

(b)

Fig. 3. SRT reference trajectory; (a) time-altitude; (b) time-velocity.

to space. The current USV vehicle configuration is a sorbital vehicle capable of a maximum speed correspoing to Mach 8. Some of the key technologies that willdemonstrated by the USV vehicle include advanced theprotection systems, Liquid Oxygen-Hydrocarbon (LOx-Hand airbreathing propulsion, intelligent and integrated hemanagement systems, and smart structures.

The USV geometric characteristics are reported inble 1. It has full span split elevons that, deflected symmrically, provide pitch control, while asymmetric deflectioprovide roll control. It also features two symmetric vertictails, with moveable rudders, for directional stability/contr

The mission profile considered in this paper correspo
to the Suborbital Re-entry Test, Fig. 3.
Fig. 4. Geometric details of the USV nose cap, dimensions in millimet

Fig. 5. Schematic of the internal structure of the USV nose cap, dimenin millimeters.

Table 2ZrB2 properties [29]

ZrB2 properties

ρ, density (kg m−3) 6000cp , specific heat (J kg−1 K−1) 628λ, thermal conductivity (W m−1 K−1) 66ε, emissivity 0.8

The USV vehicle will be released by a balloon at an atude of about 35 km. The vehicle will be powered by a sorocket motor and it will reach an altitude of about 120 kfrom which it will start the descent into the atmosphere. Tflight duration from balloon release to landing is about 35

A detailed representation of the nose cap of the USVshown in Fig. 4. The nose radius of curvature isRn = 1 cm.The initial 40 cm of the nose cap may be considered acone-sphere with an angle of inclination of 15 deg. Numcal simulations have been limited to this conical part. A Tmodel has been assumed, consisting of an initial bulk p10 cm long, followed by a shell 0.3 cm deep, Fig. 5. TTPS material is assumed to be Zirconium Diboride, whproperties are listed in Table 2 (as taken from [29]).

3. Numerical model

The axisymmetric Navier–Stokes equations have b
solved using a finite volume numerical code. Non-dissipative


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Fig. 6. Particular of the computational mesh in the fluid (a) and solidregions.

flux computations are based on the Flux Difference Split(FDS) technique, with the corresponding linearized jabian matrices obtained from Roe’s method [24]; dissipafluxes are computed using central schemes. For steadysimulations, pseudo-time integration is performed usingexplicit multi-stage Runge–Kutta scheme. For time depdent simulations, the numerical integration of the equatiin time is performed using a second-order implicit methA more detailed description of the numerical technique chacteristics may be found in [21].

The computational mesh is shown in Fig. 6. Prior toextensive computations performed in this work, a prelinary grid convergence analysis has been carried out tothe best compromise between CPU time and results aracy.

The energy equations in the fluid and the solid are cpled by the local balance between convective, conducand radiative heat fluxes:

−λ

(∂T

∂n

)w

= −λs

(∂T

∂n

)s,w

+ εσT 4w (1)

whereλs is the solid thermal conductivity.Eq. (1) states that heat transferred to TPS by convec

processes is then transferred away from the sharp leaedge by conduction in the solid material and re-radiafrom the TPS surface.

Two different approaches have been considered to cpute the time evolution of the surface temperatures andflux distribution during vehicle reentry:

(a) Fully coupled method in which the fluid-dynamic fieand the thermal field in the solid are simultaneousolved. In this case fluid and structure are coupthrough the heat flux continuity at the solid-fluid iterface, which is represented by condition (1). Thisproach is computationally expensive. However, the rof fluid-dynamic to thermal characteristic times is of torder ofα/RnV , α being the material thermal diffusiv
ity andV the flight speed. This ratio, in the present case
e

-

t

(see Fig. 5 and Table 2), is of the order of 10−6. Thissuggests, as an alternative computational procedursolve the flow field around the body and the tempeture field in the solid independently, if suitable couplischemes are used. This question is addressed in thebullet.The coupled computational approach has been validon a geometrically similar configuration (a blunted cosee [27]). The numerical results show, in this casegood agreement with the available experimental dwith differences, in terms of surface heat fluxes, lthan 10%.

(b) Uncoupled method: In this approach the vehicle tratory is discretized in time intervals with amplitudes�t . The fluid-dynamic equations and the energy eqtion in the solid are solved sequentially. Since the fludynamic characteristic times are, as previously shoseveral orders of magnitudes lower than the thercharacteristic times, the computational fluid-dynamsimulations may be performed steady per each timeterval. Once the surface heat flux is computed bysteady aerodynamic simulation at timet , the solution ofthe energy equation in the solid, in the time intervalt ,t + �t), is obtained by the following steps:

1. The convective heat transfer coefficient at timet ,h(t), is evaluated from the heat flux distributioqw(t):

h(t) = qw(t)

H0(t) − hw(t)

where H0(t) is the total specific enthalpy corresponding to the flight conditions at timet along theUSV trajectory, andhw(t) is the local specific walenthalpy.

2. The energy equation in the solid is solved intime interval (t, t + �t), assuming as boundarcondition a time-dependent surface heat fluxq(t):

q(t ) = h(t)(H0(t) − hw(t )

) − εσT 4w(t ) (2)

with t ∈ [t, t + �t].Condition (2) is obtained by taking into account tradiative heat flux at each timet .This computation provides the new surface tempature distribution,Tw(t + �t), which is then inputto the CFD algorithm to update the fluid-dynamfield and the surface heat flux,qw(t + �t), at theflight conditions corresponding to timet + �t .

This method is less expensive, from a computatiopoint of view, than the coupled approach, requiring oone steady CFD computation and an unsteady, fast, the
analysis in the solid, per time interval.


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Fig. 7. Temperature evolution in the fluid and solid regions. Coupmethod. (a)t = 30 s,z = 45 km; (b)t = 60 s,z = 20 km.

4. Results

The numerical simulations were carried out in the raof altitudes between 80 and 20 km of the reentry trajecto

The vehicle structure is assumed to be initially at a teperature of 300 K. The aerodynamic and the structure tmal analysis are performed using the “coupled” and the “coupled” methods described in Section 3.

Figs. 7(a), (b) show the fluid and solid temperature disbutions obtained with the fully coupled approach, after 3and 60 s, respectively. The corresponding altitudes alongreentry trajectory are 45 and 20 km. The temperature inthe UHTC structure gradually increases with time, achievthe maximum values in the nose region, where convecheat fluxes are larger. After 60 s, the maximum temperain the solid is about 1600 K.

Figs. 8(a), (b) show the temperature distribution insolid structure, at the same conditions as in Figs. 7, obtausing the decoupled approach and a time interval�t = 1 s.The temperature contours shown in Fig. 8 do not exhibitpreciable differences with respect to the fully coupled coputations, even if the required computational time is low

The stagnation point convective heat fluxes and tempture profiles along the trajectory are shown in Figs. 9(a),The maximum stagnation point heat flux, attained att =
55 s, is about 3.5 MW m−2, corresponding to an altitude of
(a)

(b)

Fig. 8. Temperature distribution in the solid region. Decoupled appro(a) t = 30 s,z = 45 km; (b)t = 60 s,z = 20 km.

about 22 km, while maximum temperature is reached aabout 60 s (z = 18 km). The peak temperature, correspoing to the peak heat flux, is about 1600 K; this value is ab800 K lower than the local radiative equilibrium temperat(i.e. the temperature corresponding to zero conduction insolid) and also about 300 K lower than the overall radiaequilibrium taking into account the conductive flux acrothe solid (see Fig. 9). In fact, the energy entering the surby convection accumulates in the TPS during the transphase in addition to being re-radiated into the atmosphand conducted through the solid, according to the equat

d

dt

∫V

ρcpT dt +∫S

q · ndS −∫S

σεT 4 dS = 0 (3)

whereV is the TPS volume,S the TPS surface,q is the heatflux vector andn the local surface normal.

Ideally, only in quasi-steady conditions the convectiveergy flux is distributed in the solid and re-radiated inatmosphere, so that overall radiative equilibrium is achiev

Comparison between the two computational methocoupled and decoupled, shows a satisfactory agreemWhen the decoupled approach is used with a time ofthe results are almost coincident (differences of the oof 1%) with those obtained with the fully coupled methoEven when a larger time step is utilized, temperature
convective heat fluxes differences along the trajectory do not


g the

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, ases

thethatnalameod

italare

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filedis-nceK).and

er-sults

siv-of

(a)

(b)

Fig. 9. Stagnation point heat flux (a) and temperature (b) profiles alonSRT trajectory.

exceed 10%. In the latter case the CPU time is reduceone order of magnitude.

In this case the decoupled approach works well sinceshown above, fluid-dynamic and thermal characteristic timdiffer by several orders of magnitudes. This would not becase if the TPS structure was made of a thin shell. Incase, the characteristic thermal diffusion time, proportioto the square of the shell thickness, could be of the sorder of magnitude of the fluid dynamic time and the methwould not provide accurate results [22].

The surface temperature evolution along the suborbreentry trajectory is shown in Fig. 10. The peak values
well below the maximum allowable temperatures for UHTC
Fig. 10. Surface temperature evolution.

Fig. 11. Time evolution of the stagnation point temperature for three dient values of emissivityε; and thermal conductivityλ.

materials. Att = 67 s, the temperature in the nose regis 1320 K, lower than the corresponding value att = 60 s(T = 1560 K), in agreement with the temperature proshown in Fig. 9(b). Note that the surface temperaturetribution is not uniformly decreasing, but, in correspondeof the bulk-shell junction, it increases (up to about 700This is due to the presence of the thin shell configurationits associated small thermal mass.

A number of numerical computations have been pformed, using the uncoupled approach, to assess the resensitivity to the thermal conductivities and surface emisities of the TPS material. Fig. 11 shows the time profile
the stagnation point temperatures for different emissivities


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Table 3Influence of TPS thermal conductivity and surface emissivity on maximTPS temperatures (K)

λ (W m−1 K−1) ε

0.1 0.5 0.8

10 2303 2176 210450 1676 1646 1626

100 1394 1380 1371

Fig. 12. Heat flux on body wall atZ = 22 km. Free flight.

and fixed values of thermal conductivity (λ = 1 W m−1 K−1

andλ = 10 W m−1 K−1). As expected, a lower surface emsivity increases temperature. In particular, the peak station point temperature along the reentry trajectory increfrom 2100 K, for ε = 0.8, to 2300 K, for ε = 0.1 andλ = 10 W m−1 K−1, but the sensitivity to the emissivity ilarger for smaller values of the thermal conductivity (λ = 1 W m−1 K−1 an increase of about 500 K is expectfor the same variation ofε from 0.8 to 0.1, Fig. 11).

Table 3 summarizes the results of this study. In partlar, Table 3 also confirms that, when thermal conductivitsufficiently high, the effect of surface emissivity is almonegligible. Table 3 also shows that an ideal UHTC matefor TPS applications should have a thermal conductivityhigh as, at least, 100 W m−1 K−1.

4.1. Surface catalysis in flight and wind tunnel condition

Numerical simulations have been performed at flight cdition corresponding to the altitudes of 22 km, to investigthe effects of surface catalysis on vehicle TPS heating.simulation is carried out assuming steady condition andcal radiative equilibrium, i.e. neglecting the conductionthe solid.

Both fully and non-catalytic wall conditions have besimulated. Fig. 12 shows the surface heat flux distributioThe result corresponding to non-catalytic and fully cataly
computations are almost coincident, since the chemical re-
Table 4Comparison of free flight and wind tunnel conditions for stagnation pheat flux duplication

Z = 22 (km) Free flight Wind tunne

p (Pa) 4000 1980T (K) 217 388M 8.02 7p02 (Pa) 335720 125851H0 (MJ kg−1) 3 5O 0 0.018N 0 0

actions are “frozen” in the shock layer, due to the relativsmall shock stand-off distance, typical of geometries witsmall degree of bluntness.

However, surface catalysis may play a role for TPS ting in plasma wind tunnel, even on sharp geometries, siin this case, the arc-heated hypersonic stream is partdissociated in the test chamber. Therefore, numerical slations have been performed to study catalycity effects onUSV nose in the SCIROCCO plasma wind tunnel availaat Italian Aerospace Research Center. The operating wtunnel conditions have been selected in such a way toplicate in full scale the free flight stagnation point heat fli.e. assuming that the product of total specific enthalpythe square root of the stagnation point pressure is the scompared to free flight conditions, according to the Zoempirical correlation formula for the heat flux [32]:

(H0 − hw) · √p02 = cost (4)

wherehw is the wall specific enthalpy.Values ofH0 andp02 and the corresponding flow cha

acteristics in the test chamber of the wind tunnel have bselected from the SCIROCCO plasma wind tunnel enve[4]. Table 4 shows a comparison of the free flight conditioand the corresponding wind tunnel values. By considethe stagnation pressures and total enthalpies reported ible 4, and the values ofhw in the stagnation point obtainefrom computations, it is easily verified that the producthas almost the same value (about 255 MJ kg−1 Pa1/2) forboth the free flight and the selected wind tunnel conditio

Fig. 13 shows the surface heat flux distributions. Tstagnation point heat flux values are shown in Table 5.expected, about 10% higher heat fluxes occur at the station point region and far away from the nose tip, althoughspecific total enthalpy is not very high (less than 5 MJ kg−1)and consequently the oxygen is only partially dissociatethe PWT test chamber (see Table 4).

4.2. Laminar-turbulent transition effects

The boundary layer laminar-turbulent transition in hypsonic flows is dependent upon a number of parameters. Mimportant are Reynolds and Mach number effects. Starfrom Mach numbers of about 5, it is known from expe
mental data and linear stability theory that compressibility


to

.

cals is

essnetedowhenbe-the

lent

sedmallop-froming

nor-pro-e is

ore

sti-an

thelo-

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tionnd-onsi-facentum0 Kthe

si-ind

ue toe

Fig. 13. Heat fluxes on body wall at a PWT condition equivalentZ = 22 km (Table 4).

Table 5Comparison between free flight and wind tunnel results

Tmax (K) qmax (MW m−2)

Free flight (22 km) 2209 1.07PWT 2090 0.92Non-catalytic wallPWT 2211 1.15Fully catalytic wall

Fig. 14. Entropy layer on a sphere-cone geometry in supersonic flow

has a stabilizing effect on the transition. A semi-empiriformula that takes into account this effect for sharp conegiven by Anderson [1]:

log10Rextr = 6.421· e1.209·10−4·M2.641

e (5)

where Rextr = ρeUextr/ae andMe = Ue/ae .It is also well known that a degree of nose bluntn

modifies the location of the transition point along the cosurface [2]. This is an effect of the entropy layer associawith the bow shock wave (see Fig. 14), that alters the flproperties at the boundary layer edge. In particular, wthe ratio of the nose radius to the cone base radius islow 0.5–0.6, nose bluntness has a stabilizing effect on
laminar boundary layer, moving the transition location fur-
Fig. 15. Surface temperature distribution: effect of the laminar-turbutransition.

ther downstream. However, correlation (5) may still be uwith confidence in the present case, due to the very sdegree of nose bluntness. Similar to Ref. [13], the flow prerties at the boundary layer edge have been extractedthe CFD laminar results using an “ad-hoc” post-processprocedure. The algorithm computes the body surfacemal at each computational point on the surface andceeds along each normal until the boundary layer edgfound, i.e. when two consecutive values of U differ no mthan 1%.

Laminar-turbulence transition effects have been invegated considering the flight conditions corresponding toaltitude of 22 km along the reentry trajectory, wherelargest thermal load occurs. Using Eq. (5), the transitioncation is estimated to be atx = 0.28 m. Therefore a furtheCFD simulation has been carried out solving the lamiNavier–Stokes equations in the region before the transpoint and the Reynolds Averaged Navier–Stokes equatwith standardk − ε turbulent formulation forx > 0.28 m,taking into account the turbulence effects after the transilocation. The surface temperature distributions correspoing to overall radiative equilibrium conditions in the twcases are shown in Fig. 15. The laminar-turbulent tration has a non-negligible effect on surface heating: surtemperature, due to the augmented exchange of momeand heat transfer, increases from about 1400 K to 190at x = 0.36 m, attaining values comparable to those instagnation region.

It is important to underline that laminar-turbulent trantion cannot be reproduced at the SCIROCCO plasma wtunnel conditions discussed in the previous paragraph, dthe very low Reynolds number (Re= 83200, based on con
length) of the plasma stream.


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5. Conclusions

The aerodynamic performance of future hypersonic vcles could be greatly improved with sharp leading edgesharp fuselage tips. Numerical simulations of the aerotmal environment around the nose cap of the USV have bcarried out, to show the importance of performing timdependent simulations, in order to correctly predict the tperature distribution in the thermal protection material.

The results of the analysis show that a desirable TPSterial should have relatively high thermal conductivity, of torder of 100 W m−1 K−1, and surface emissivity (more tha0.5).

Ultra High Temperature Ceramics, which can operattemperatures up to 3000 K and which exhibit desirable tmal conductivities and emissivities, seem particularly wsuited for use as massive leading edges of sharp wingsfuselage noses, protecting the rest of the structures whevehicle flies at relatively low angles of attack (< 10 deg).

However, the “boundary layer thermal protection” [17]only possible if laminar conditions prevail. Near the laminturbulent transition, heat fluxes comparable to the stagnapoint heat flux occur far from the stagnation region, resultin a temperature increase of about 500 K, and reducingadvantages of the laminar boundary layer thermal protec

In this case, the simulation of the flight conditionsplasma wind tunnel poses serious problems, not onlycause surface catalycity becomes important, but also beclaminar-turbulent transition cannot be reproduced

References

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Aerothermodynamic Study of UHTC based TPS

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