Aeroheating Environments for a Mars Smart Lander

14 Pages·2002·1.2 MB·English

by  Edquist, Karl T.

AIAA 2002-4505
Aeroheating Environments for a
Mars Smart Lander
Karl T. Edquist, Derek S. Liechty, Brian R. Hollis,
and Stephen J. Alter
NASA Langley Research Center
Hampton, VA
Mark P. Loomis
NASA Ames Research Center
Moffett Field, CA
AIAA Atmospheric Flight Mechanics
Conference & Exhibit
5-8 August, 2002
Monterey, California
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AIAA 2002-4505
AEROHEAT1NG ENVIRONMENTS FOR A
MARS SMART LANDER
Karl T. Edquist*, Derek S. Liechty t, Brian R. Hollis;, and Stephen J. Alter _
NASA Langley Research Center
Hampton, VA
Mark P. Loomis ¶
NASA Ames Research Center
Moffett Field, CA
ABSTRACT
A proposed Mars Smart Lander is designed to reach flae surface via lifting-body atmospheric entry (t_ = 16
deg) to within 10 km of the target site. CFD predictions of the forebody aeroheating environments are given for a
direct entry from a 2005 launch. The solutions were obtained using an 8-species gas in thermal and chemical non-
equilibrium with a radiative-equilibrium wall temperature boundary condition. Select wind tunnel data are
presented from tests at NASA Langley Research Center. Turbulence effects are included to account for both smooth
body transition and turbulence due to heatshield penetrations. Natural transition is based on a momentum-flaickness
Reynolds number value of 200. The effects of heatshield penetrations on turbulence are estimated from wind tunnel
tests of various cavity sizes and locations. Both natural transition and heatshield penetrations are predicted to cause
turbulence prior to the nominal trajectory peak heating time. Laminar and turbulent CFD predictions along the
trajectory are used to estimate heat rates and loads. The predicted peak turbulent heat rate of 63 W/cm: on the
heatshield leeward flank is 70% higher than the laminar peak. The maximum integrated heat load for a fully
turbulent heat pulse is 38% higher flaan flae laminar load on the heatshield nose. The predicted aeroheating
environments wifla uncertainty factors will be used to design athermal protection system.
NOMENCLATURE H velocity parallel to surface (m/s)
V velocity (km/s)
CD drag coefficient, D/½p_V2S W full-scale heatshield cavity diameter (in)
D heatshield diameter (m) )C radial coordinate (m)
h altitude (kin) y+ non-dimensional boundary layer coordinate
L/D lift-to-drag ratio trim angle of attack (deg)
m aeroshell mass (kg) ballistic coefficient, m/CDS (kg/m 2)
p pressure (N/m2) 6 boundary layer thickness (in)
Q heat load (J/cm2) 0 circumferential location on heatshield
q heat rate (W/cm 2) measured clockwise from leeside (deg)
R heatshield radius (m) Y flight path angle (deg)
Rn nose radius (m) viscosity (kg/m2-s)
R_ shoulder radius (m) 0 momentum flaickness (m)
Re Reynolds number
Oc cone half angle (deg)
r radial coordinate (m)
P density (kg/m s)
S reference area, _!)2/4 (m2)
_mox maximum grid stretching in the surface normal
s running length (m)
coordinate direction
T temperature (K)
TPS Thermal Protection System
Subscripts
t time from atmospheric interface (s)
D heatshield diameter
*Vehicle Analysis Branch, Senior Member, AIAA.
Aerothermodynamics Branch. e boundary layer edge
*Aerothermodynamics Branch, Senior Member, AIAA. FR Fay-Riddell value
§Aerothermodynamics Branch, Senior Member, AIAA. lain laminar
¶Reacting Flow- Environments Branch, Member, AIAA.
turb turbulent
Copyright © 2001 by the American Institute of Aeronautics and w wall
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copyright owner.
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INTRODUCTION forebody aeroheating environments for flae proposed
MSL mission using a combination of CFD calculations
First generation Mars landers such as Viking and and wind tunnel data.
Paflafinder successfully used entry trajectories that
provided landing accuracy within hundreds of Heatshield Geometry
kilometers of their targets. Second generation landers
are characterized by their ability for precise landing (< The heatshield geometry used for flae aeroheating
10 kin) near areas that are of particular scientific predictions is a 4.05-m diameter, 70-deg half-angle
interest 1. The Mars Smart Lander (MSL) is a proposed sphere-cone shown in Figure 2. Table 1 summarizes
mission that is designed to achieve precise landing the heatshield dimensions. The forebody geometry is
accuracy at sites flaat were previously unreachable. similar to flaat of Paflafinder, but much larger (4.05 m
vs. 2.65 m). A stable trim angle of 16 deg is required
Figure 1shows aproposed MSL configuration in for hypersonic L/D= 0.22-0.25 and is achieved with a
which flae cruise stage is attached to the lander pallet radial center-of-gravity offset.
through flae heatshield. The rover occupies much of the
Rs
interior volume and requires a larger backshell flaan was
used for Viking and Paflafinder. The aeroshell is
designed to separate from flae cruise stage prior to
atmospheric interface and flaen deliver the lander to flae
Martian surface.
Backshell
Parachute
/
Rover D
/
Lander
Pallet
Figure 2. Heatshield Geometry
Table 1. Heatshield Dimensions
Heatshield
Cruise jj.J "_ , Descent Parameter Dimension
Stage Engines D (m) 4.05
Figure 1. MSL Cruise Configuration Oc(deg) 70
Rn (in) 0.9854
A guided lifting-body entry is proposed for MSL R_ (m) 0.0988
to achieve 10-kin landing accuracy. The entry vehicle
lift is used to mitigate uncertainties in predicted entry Heatshield Penetrations
states, atmospheric properties, aerodynamics, etc. that
would otherwise contribute to large landing footprints. If the entry vehicle is attached to flae cruise stage
A hypersonic lift-to-drag ratio (L/D) of 0.22-0.25 and through flae heatshield, cavities will remain after cruise
ballistic coefficient (tim) less than 120 kg/m: meet flae stage separation. Figure 3 shows six equally spaced
delivery requirements for a direct entry from the 2005 cavities of diameter W at radial location r that are left
launch opportunity. behind when the bolts connecting flae pallet and cruise
ring are severed. The cavity depth is expected to be
The aeroshell contents will require protection smaller flaan the diameter. The cavities could
from exposure to significant aeroflaermal environments significantly augment flae heating environment in the
during entry. Knowledge of the expected heating is vicinity of the penetrations and downstream. A
necessary for proper design of the thermal protection combination of CFD and wind tunnel test data was used
system (TPS). This paper summarizes flae predicted to investigate flae penetration effects on aeroheating.
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ANALYSIS
Computational Approach
CFD calculations at flight conditions were
performed using the Langley Aerothermodynamic
Upwind Relaxation Algorithm (LAURA) 3 and _General
Aerodynamics Simulation Program (GASP) 4 CFD
codes. Both codes were exercised using an 8-species
Mars gas (CO2, CO, N2, O2, NO, C, N, O) in thermal
and chemical non-equilibrium with a radiative-
equilibrium wall temperature boundary condition. The
Park-945 reaction rates were used for the 8-species
chemistry model in both codes. A "super-catalytic"
wall boundary condition was used in which the species
mass fractions for CO2 and N2 are fixed at their
Figure 3. Lander Pallet/Cruise Stage Attachment freestream values of 0.97 and 0.03, respectively. This
Points through Heatshield assumption gives the most conservative heating levels.
LAURA was used for a series of aeroheating
Nominal Entry Trajectory predictions along the nominal trajectory and GASP was
used to support those results. Only convective heating
Aeroheating environments analysis requires is predicted because the contribution t_om radiation
knowledge of the entry trajectory and associated heat should be negligible.
pulse. Trajectories for the 2005 launch opportunity
were calculated using the Program to Optimize LAURA CFD Code
Simulated Trajectories (POST) 2. A total of 2000 three
degree-of-freedom Monte-Carlo simulations was LAURA was developed at NASA Langley
performed to investigate dispersions in entry states, Research Center (LaRC) and has been used previously
atmospheric properties, and aerodynamics, among to predict the aeroheating environments for various
others. The nominal, or expected, entry trajectory was Mars projects 6'7. The code uses a finite-volume
selected for initial aeroheating analysis. A velocity of approach to solve the inviscid, thin-layer Navier-Stokes
5.56 km/s and flight path angle of -12.5 deg at (TLNS), or full Navier-Stokes flowfield equations. The
atmospheric interface produce the velocity-altitude map TLNS option was used for all LAURA calculations
shown in Figure 4. Peak stagnation point heating presented here. Roe's averaging 8 is used for the
occurs at an altitude of 37.1 km and a velocity of 4.92 inviscid fluxes with second-order corrections using
km/s. CFD predictions of the heatshield aeroheating Yee's symmetric total variation diminishing (TVD)
environments are shown for the nominal trajectory. scheme 9. Turbulent LAURA solutions were obtained
using the Baldwin-Lomax algebraic model 1°'11. On a
Atmospheric !nterface simple sphere-cone geometry, the Baldwin-Lomax
V = &56 km/$ -.
_,_.12saeg _- model is believed to give reasonable results. A user-
specified transition location is required to run the model
and was specified to give fully turbulent results for
conservatism.
E A built-in LAURA grid alignment capability
Peak Heating allows mesh adaptation to the boundary layer and bow
V_ = 4.92 kmts shock according to user-defined parameters. Proper
50
Parachute _ J cell spacing at the wall is important for heating
Deployment _i_" calculations and is controlled in LAURA by the wall
cell Reynolds number, which is defined as:
0i 1 2 3 4 5 6 paArl
v,. (kmts) (1)
Figure 4. Nominal Entry Trajectory for the 2005
Launch Opportunity
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where
p is the density, a is the speed of sound, At/is
the cell height, and bt is the viscosity. The author's
experience is that reliable laminar heating predictions
can be achieved using Re,; = 10. A Re,; value of 2 was
used for turbulent calculations to give a y+ value of
order 0.1, where y+ is defined as:
(2) /.::L:';,I ¸; i: •_i
I :/-,i _?'?. 2,?_
The quantities r/and _',_are the surface normal distance ',5.Y
and shear stress, respectively.
GASP CFD Code
GASP has been used extensively at NASA Ames %%<;,,_i.:ii:iJ)i
Research Center in TPS sizing applications for both
planetary 12'13 and tzansatmospheric 14 entry vehicles.
The GASP solutions were run with models that are
Figure 5. Heatshield Surface and Symmetry Plane
similar, if not identical, to those used in LAURA. The
Grids Used for LAURA Solutions (Coarsened)
transport properties are calculated with Wilke's 15
mixing rule with curve fits for the species viscosities CFD Solution Points
given by Blottner, et al. 16and Eucken' S17 correlation for
thermal conductivity. Binary diffusion was used with a Detailed aeroheating environments prediction at
constant Schmidt number of 0.7, which indicates the flight conditions requires high-fidelity CFD calculations
ratio of momentum and mass diffusivities. The Van
at multiple points along the design trajectory. Figure 6
Leer 18flux splitting scheme with the min-mod limiter is shows 12 points on the 2005 nominal trajectory at
used to calculate the inviscid fluxes and a central- which LAURA solutions were obtained. Some points
difference approximation is used for the viscous fluxes. were identified to analyze specific milestones, such as
For turbulent calculations, the Baldwin-Lomax model peaks in stagnation point heating (t = 103 s), dynamic
was used with acompressibility correction. pressure (t = 119 s), and Reynolds number (t = 130 s).
The remaining points were selected to fill out the heat
Computational Grid pulse with sufficient resolution. The heat rate shown in
Figure 6 is based on the Fay-Riddell formula 22and is
A 7-block, singularity-free, structured volume not corrected for angle-of-attack. The entire heat pulse
grid was used for the LAURA solutions. The lasts about five minutes, with most of the heating
heatshield nose was meshed to avoid a singularity pole occurring at t = 45-190 s.
boundary because it can introduce discontinuities in the
flowfield solution. Figure 5 shows the surface and I
...................... Dynamic PP_ssure j
symmetry plane grid distributions coarsened by a factor ........... R_ 1
of two in each direction. Only half of the heatshield is
modeled due to symmetry in the pitch plane. The grid
was built using GridGen 19to construct the topology and 40
surface distribution, 3DGRAPE/AL 2° to generate the
volume grid, and the Volume Grid Manipulator N ao ;i ",,. 40Q0
(VGM) 21to enhance grid quality and accurately impose laE_
boundary conditions along block interfaces.
A total of 3280 surface and 64 normal cells was
used to resolve the flowfield. The circumferential mesh
distribution is equally spaced at 3-deg increments for a 0 ...aOLOE_O0
:3 100 200 300
t(st
total of 60 cells. A relatively fine streamwise grid
distribution is used on the nose and shoulder to Figure 6. CFD Solution Points along Nominal
accurately reproduce the surface geometry and capture Trajectory
steep flowfield gradients in those regions.
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Table2 summarizfelasefreestxeacmonditions
dimensional shapes can be collected with less cost and
forflaeCFDsolutionsA.llflightcasewsererunat
o_= lead-time requirements than with discrete gauges.
16 deg, which is required for a hypersonic L/D = 0.22-
0.25. A test model with cavities is shown in Figure 7.
During a tunnel run, the phosphor-coated model
Table 2. Freestream Conditions for CFD Solutions fluoresces under ultraviolet light according to surface
temperature. Model temperature before and after arun
t (s) h ¢:m) is digitized and reduced to a heat rate using one-
45 80.6 5.59 8.31 x 10-7 108.8 dimensional heat conduction theory. The total
61 66.2 5.59 8.26 x 10-0 116.1 experimental uncertainty is estimated to be +/- 13% for
the conical flank. This estimate takes into account
70 58.9 5.58 2.59 x 10-5 120.2
uncertainties in freestxeam conditions, fluorescent
83 49.2 5.48 1.02 x 10-4 133.4
intensity, data extraction, and model conduction.
93 42.6 5.28 2.26 x 10-4 147.5
103 37.1 4.92 4.15 x 10-4 159.3
112 33.3 4.47 6.20 x 10-4 167.6
119 31.1 4.07 7.82 x 10-4 170.5
130 28.8 3.47 9.99 x 10-4 173.6
140 27.5 2.98 1.14x 10-3 175.3
155 26.7 2.42 1.24 x 10-_ 176.5
190 25.6 1.64 1.39 x 10-3 178.1
Experimental Approach
Tests were conducted in NASA LaRC's 20-Inch
Mach 6 Air Tunnel to investigate the aeroheating
characteristics of various proposed MSL
configurations. The test objectives were to compare the
heating environments on different geometries and
determine the effects of heatshield penetrations on Figure 7. Phosphor-Coated Tunnel Model with Six
turbulent transition and heating augmentation. Table 3 Cavities (Black Dots are Fiducial Marks)
lists the pertinent test parameters. A range of Reynolds
numbers (Re=), angles-of-attack, and cavity diameters RESULTS AND DISCUSSION
and locations was investigated. References 23-25
contain detailed discussions of the aeroheating test Laminar and turbulent CFD flight predictions are
objectives and results. Reference 25 also includes presented for the nominal trajectory from a 2005
heating data on asymmetric heatshield shapes. launch. The laminar results were used for the transition
analysis and the turbulent cases provide estimates of
Table 3. Aeroheating Test Parameters turbulent heating augmentation.
Parameter Value Transition to Turbulence
Model Diameter (in) 5 (3.1% scale)
Mach number 6 Transition via two mechanisms was analyzed for
the MSL heatshield with penetrations: natural smooth
Reynolds number 2.6 -7.3 x 106/fi
body transition due to boundary layer instabilities and
Angle of attack (deg) 0, 11, 16, 20
tripped turbulence caused by heatshield cavities.
Cavity Diameter, W (in) 1, 2.2, 3 (full scale) Surface roughness and ablation product effects on
Cavity Depth (in) O.3W transition have not yet been analyzed and are a function
Cavity Location (r/R) 0.41, 0.7 of the TPS material.
Global surface heating distributions were Smooth Body Transition
obtained on ceramic models using the two-color,
relative-intensity, phosphor thermography method 26'27. The prediction of natural turbulent transition in
Phosphor tkermography is the standard method at the hypersonic regime is an ongoing effort 2s'29. The
LaRC for obtaining global surface heating in the influence of boundary layer instabilities on turbulence
center's hypersonic tunnels. Heating on complex three- is not well known and is a function of many parameters
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thatareuniquetoa givenconfiguratioanndflow
A conservative transition criterion of Reo > 200
conditionsA. knownfact,whichhasbeenobserved
was used for MSL and is based on previous experience
repeatedilsyt,hatturbulencoeccurmsorereadilywith
with hypersonic blunt bodies 3°. The effect of the trim
increasinRgeynoldnsumberO. necommomnethoodf
angle on the boundary layer is shown in Figure 9. Non-
predictintgransitionw,hichis basedona constant
dimensional surface pressure is plotted with Reo = 200
momentum-thickRneeysnsoldnsumbe(rReow),asused
and boundary layer edge Mach number (Me 1) at the
toanalyzteurbulentrtansitiofnortheMSLheating
nominal trajectory peak heating point. Based on the
environmentTsh.edefinitioonf
Reo is based on the transition criterion, turbulence is predicted for most of
boundary layer edge properties and momentum the leeward side of the heatshield, where running length
thickness (0), which is a measure of the momentum and edge velocity are largest. In contrast, turbulence is
deficit due to the boundary layer: not predicted for the windward side of the heatshield.
A subsonic boundary layer exists on most of the
heatshield and may promote transition in the presence
peHe 0
Reo - (3) of surface roughness and/or mass addition from ablation
products.
8= _ Pu II-u--Idrl (4)
opeu_ \ u_)
The boundary layer edge is defined as the location
where total enthalpy is 99.5% of the freestream value.
The trim angle of 16 deg has strong implications
for the stagnation point location and propensity for
turbulent transition. The streamlines in the symmetry
plane and near the surface are plotted with non-
dimensional pressure at the nominal trajectory peak
heating point in Figure 8. At _ = 16 deg, the stagnation
point moves off the heatshield nose and onto the
bottom (windward) flank. This stagnation point
location results in ashort boundary layer running length
on the windward side and, more importantly, a long
running length on the top (leeward) side. Boundary
layer instabilities are more likely to occur when a long
Figure 9. LAURA Surface Pressure, Reo= 200, and
running length exists.
Me = 1 at Nominal Peak Heating (G= 16 deg)
Figure 10 shows LAURA laminar symmetry
plane Reo distributions at select times along the nominal
trajectory. The peak Reovalues are plotted as a
Leeward function of time with freestream dynamic pressure in
Figure 11. The Reo values are highest on the leeward
p/p,V_: flank where the running length and edge velocity are
:::: 00.6912 large; the lowest magnitudes are located on the
069
058 windward side near the stagnation point. Note that Reo
16° 047
o._6 closely follows dynamic pressure and their peak values
Stagnation 000..021243 occur at the same time. Transition is predicted slightly
Point before t = 83 s on the leeward side, but never on the
windside. Since transition is predicted before the peak
Windward
heating time, turbulent augmentation will be shown to
have a significant effect on the peak heat rate and
integrated heat load delivered to the heatshield.
Figure 8. LAURA Symmetry Plane and Surface
Pressure and Streamlines at Nominal Peak Heating
(I_= 16 deg)
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Heatshield Penetrations
The existence of heatshield cavities may also
cause turbulence and augment heating. Heatshield
4¢0 penetrations are being used for flae Genesis Sample
Return Capsule (SRC); CFD and wind turmel data were
used to predict flae influence of the cavities on
aeroheating 31. The same approach is used for MSL in
2°°- ........... which the penetration effects are reduced to a function
of the penetration size (/4/), local boundary layer
thickness (o_, and Reo. CFD and tunnel data were used
100_ i!i to predict combinations of penetration configurations
and flowfield conditions that result in a tripped
turbulent boundary layer and augmented heating. Table
x(m) 4 lists the various cavity sizes and locations tested on
3.1% scale models in flae LaRC 20-Inch Mach 6 Air
Figure 10. LAURA Symmetry Plane Reo along Tunnel.
Nominal Trajectory ((_= ]6 deg)
Table 4. Aeroheating Test Heatshield Cavity
Parameters
Cavity W (in, full scale) r/R
1 3.0 0.7
2 2.2 0.7
3 1.5 0.7
4 3.0 0.41
5 2.2 0.41
6 1.5 0.41
In general, larger cavities at higher freestream
Reynolds numbers (Re=), and flaus higher Re_ are more
likely to trip flae boundary layer. The effect of Re= on
the penetration's ability to cause turbulence is shown in
Figure 12. The family of curves represent heat rate
profiles behind a leeside 2.2-in penetration (on a4.05 m
Figure 11. LAURA Peak Reo and Freestream
Dynamic Pressure along Nominal Trajectory aeroshell) for various Re= at _= 16 deg. The auxiliary
(_ = 16 deg) figures show the cavity locations and phosphor
thermography results at the highest Re= (7.3 x 106/ft);
all but cavity number 2 increases the heating in wedge-
It should be noted flaat flae Reo transition criterion
shaped regions behind the penetrations.
was originally developed with engineering boundary
layer codes. It has been observed that flae boundary
layer edge properties predicted by CFD and For a lowRe= (3.0 x 106/fi), the heating behind
engineering codes can differ significantly. The result is cavity 5 is the same as the laminar smooth body level.
that CFD codes often predict higher Reo values (by a With increasing Re_, flae cavity is more effective in
factor of 2) than do engineering codes, and thus have causing turbulence and augmenting flae downstream
some built-in conservatism. However, even if the MSL heating. At the highest Re= (7.3 x 106/fi), flae heating
Reo transition criterion was doubled to 400 to account asymptotically reaches a non-dimensional turbulent
for conservative CFD results, transition would still be level of qt_b/qFR = 1.08, compared to a peak laminar
predicted before the peak heating time. Also, a nose value of approximately q1,,/qFR = 0.73. Thus, the
Reovalue of 200 seems reasonable at this time since laminar heat rate is increased by approximately 48%
turbulence due to surface roughness and mass addition due to cavity-induced turbulence. If similar heating
from ablation products have not yet been considered. augmentation exists in flight, the TPS requirements can
increase significantly.
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_25 predictions should be made using turbulent solutions by
...........................S...m...o..o..th Body; Re_ =3,0 x10'_Ift
..............................R..e =3:0x10_i ft at least t - 83 s on the nominal trajectory to properly
_L Re_ = &lxiO_i ft ;.,
estimate the heating envlrorLments.
..........;..:..;..................Re, = 7,3 xt0 ft ......... ,,..... :_;
,_ t=119 (Peak Dynamic Pressure)
300 t=t03 (Peak Heating) a Laminar
Local Disturbance
\..
_, _, O Transitional
Turbulent
:_ -.-.-.-.-.-.-.-.-.-.-.-.-L.-a.-m.-i-nar Boun_lary
_-- _ _ T_lfbtllentt_oulld_ry
t=_93 :, _ _ 2005 N_minal(W=3_ dR=0.41,_=01
200
0;25 [ a
cav_
O 0i25 0.5 0:';'5 1 100
dR
Niiiiiiiiii;i;;_?!;_:_ _i;i;;ii;iii;i_,_
r:._,,.,,Z2y _{;"
%, .::;::."
Figure 12. Heat Rate Data for Symmetry Plane
%._;::":ii!!!_ii:ii:_;:i_
Cavity #5 (¢_= 0) at Various Re. (t_= 16 deg)
The transition map in Figure 13 assembles all
combinations of cavity size and location, Re_, and o_
from the aeroheating tests into a comprehensive plot J
showing the penetration effects on turbulent transition.
The plot of Reo vs. lg/c_ is divided into regions of
laminar, local disturbance, transitional, and turbulent
heating levels based on the turmel data. The legend Figure 13. Heatshield Penetrations Transition Map
shows representative heating images for the different and Legend
data groups. The Reo and c_values for each point were
computed with LAURA at tunnel conditions assuming It should be noted that using tunnel data to
perfect gas and a constant wall temperature of 300 K. predict transition at flight conditions is a conservative
Some overlap exists between adjacent data groups and approach. Turmel noise can accelerate the onset of
may be due to the fact that data from all angles-of- turbulence, whereas freestream disturbances at flight
attack (0-20 deg) are placed on the same plot. Curve conditions are generally small. The implications of
fits of the data were used to delineate the regions of the turbulence on the heating environments are shown next
map and predict the penetration effects in flight. with flight predictions on the nominal trajectory.
Conservative design practice would mandate using the
laminar boundary curve as the transition indicator. See Flight Predictions
Reference 20 for a discussion of the transition map
development. Two aeroheating quantities are of particular
importance for TPS design. First, the selection of
Values of Reo and V//c_were calculated for a 3-in candidate TPS materials is limited by the peak heat rate
leeside penetration (r/R = 0.41, 0= 0) along the 2005 encountered during entry. Second, the TPS thickness
nominal trajectory. Based on these calculations, it is (and mass) is based on heat rate integrated over the
estimated that the penetration would trip the boundary entire heat pulse, or total heat load. The effects of
layer to full turbulence between t = 83 s and t = 93 s, turbulence on heat rates and loads were estimated with
which is well before the peak heating point at t = 103 s. laminar and turbulent solutions along the nominal
The smooth body transition analysis also indicated a trajectory.
transition time near t = 83 s. Therefore, flight
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Baldwin-Lomtauxrbulenstolutionisn anon-
_or ,, qsoo
equilibriuMmarsatmosphehraevelimitationsF.irsta, i . ,_i t
...... <2-........ Lammarq_, ,_ _ l
user-suppltireadnsitiolnocationisrequiretdorunthe
6Or" : R .::: z _ 3
modela,ndthusissomewhaarttificial.Secontdh,e
turbulenhteatratespredictedduringflightarehigh sog ; )t
enougthocausTePSsurfacreecessioanndpyrolysis,
neitherof whichis factoredintothe solutions.
Nevertheletshset,urbulenstolutionasrebelievetdobe
appropriateclyonservativaendareshownto be -'_ ao- J' _ / :;22o0
consistewnitththeturbulentrtansitioannalysis.
Laminar vs. Turbulent Peak Heatin_
Laminar (LAURA) and turbulent (LAURA and
GASP) solutions were obtained at the nominal -2 _1 0 _ 2
x(m)
trajectory peak heating time to compare heating levels
and validate results. The transition location for the Figure 15. LAURA Symmetry Plane Laminar and
Baldwin-Lomax solutions was specified at the Turbulent Heat Rates, and Reo at Nominal Peak
stagnation point to give fully turbulent heating levels. Heating (O_= 16 deg)
A comparison of LAURA laminar and turbulent LAURA and GASP turbulent solutions are
heat rates is shown in Figure 14 with the Reo = 200 compared in Figure 16. The GASP solution was run on
level overlaid on the contours. Figure 15 shows a line a different grid than was used with LAURA. Overall,
cut of the data along the symmetry plane. The the codes predict heat rates that are within the expected
Baldwin-Lomax solution clearly indicates that uncertainty of a CFD flight prediction. The important
turbulence is most effective in augmenting the leeward result is that both codes predict the same peak turbulent
heat rate near 63 W/cm:. Differences between LAURA
side heating. The highest laminar heat rate near 37
W/cm 2occurs on the nose and windward shoulder, and and GASP near the nose are likely due in part to the
the windward side heating is higher than it is on the models used to bridge the laminar and turbulent regions
leeward side. LAURA predicts a peak turbulent heat of the flowfield. GASP mimics immediate transition
rate near 63 W/cm 2on the leeward flank, or about 70% beginning at the user-specified location, whereas
LAURA uses the Dhawan-Narashima 32model to blend
higher than the laminar peak. Some turbulent
augmentation is predicted on the windward side, but the laminar and turbulent regions. Thus, LAURA does
much less than on the leeward side. This result is not show a rapid rise in heating on the nose like the
consistent with the Reo values in those regions. GASP solution indicates. More importantly,
downstream turbulent heat rates agree very well.
........•.._........ LAURA
/_ ,,- [
sc _ -- ;;I i_
: . ,..
A
"E 4c
sc -;. Windward ;i_ Leewa El rii
_2 0 _ 2
s(m)
Figure 14. LAURA Laminar and Turbulent Heat
Figure 16. LAURA and GASP Symmetry Plane
Rates, and Reo = 200 Contour at Nominal Peak
Turbulent Heat Rates at Nominal Peak Heating
Heating (t_= 16 deg)
(t_= 16 deg)
9
American Institute of Aeronautics and Astronautics