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August 20, 2003 1
Tutorial
Aerothermal Analysis and Design
Overview
U.S. Army AMRDECDr. Gerald Russell
Naval Air Warfare CenterRick Burnes
ITT Industries Aerotherm Forrest Strobel Dr. Al Murray
NASA Thermal and FluidsAnalysis
Workshop 2003
August 20, 2003 2
Outline
Process Overview Aerothermal Analysis & Design Process
System design requirements Aerothermal boundary conditions Candidate material technologies Analytic model development Preliminary flight design Ground aerothermal testing Flight design/verification and validation
Analysis and Design Methodologies Summary
August 20, 2003 3
Process Overview Define system requirements
Environments Material technology (TRL)
Airframe, nozzle, fins/leading edges, radomes/IR windows Internal components
Electronics, propellants, fin root bearings, seals Weight/Cost
Conduct component level & system analysis & design (shape, material, trajectories)
Conduct ground test and evaluation to validate component design models
Perform flight design Instrument flight system to further validate analytic
models
August 20, 2003 4
Aerothermal Analysis & Design Process
Select CandidateMaterial Technology
Ground TestFacilities
Flight Test withInstrumentation and
Validate Design Models
environmentsdurability, geometry
volume/weight
Propertiesapplication, weight,Volume, TRL, cost
producibility
Selection CriteriaHeat flux, shear, pressure,
total temperature, run time,test cell, cost, availability
screening testsperformance tests
Flight design usingrefine models
material technologyapplicability
phenomena of interestdefine facilities
Iteration throughoutprocess may be necessary
0.09 Char
0.14 Pyrolysis
0.2 Virgin
0.43 TotalThickness
A r e a 3
M e a n 7 1 8 . 0
A r e a 3
M e a n 7 1 8 . 0
A r e a 2M e a n 7 2 6 . 0
A r e a 2M e a n 7 2 6 . 0
A r e a 1
M e a n 7 7 8 . 9
A r e a 1
M e a n 7 7 8 . 9
A r e a 6M e a n 8 4 2 . 9
A r e a 6M e a n 8 4 2 . 9
A r e a 5
M e a n 8 5 4 . 1
A r e a 5
M e a n 8 5 4 . 1
A r e a 4M e a n 8 8 5 . 9
A r e a 4M e a n 8 8 5 . 9
*>1,881F
*
August 20, 2003 5
System Design Requirements
Reusable/single use Fabrication/manufacturing cost Schedule (TRL requirement) System integration Aerothermal environment (thermal, pressure,
surface reactions/catalysis) Rain/sand requirements Flight time Storage/transportation/environmental extremes
August 20, 2003 6
System Design Requirements Airframe/structure operating temperature
Metallics 2000F (1367 K) Ultra High Temperature Ceramics (UHTC) >3000F (1922 K) Composites (anisotropic properties)
Graphite epoxy < 350F (450 K) Cyanate ester (PT-30) < 550F (561 K) Pthalonitrile < 1100F (867 K)
Leading edge/control surface Durability (rain, cyclic heating, reusable) Thermal expansion and insulative ability Strength at temperature
August 20, 2003 7
System Design Requirements
Geometries of interest Shape (stagnation, conical, flat, wedge, leading edge,
incidence angle) Substructure (operational temperature limits) Weight/volume constraints
Aerothermal environment Trajectories (velocity, altitude, angle of attack,time) Flight geometries Recovery enthalpy/temperature Local aerodynamic shear & pressure Local heat flux (transient/integrated) Ionization/plasma potential
August 20, 2003 8
Aerothermal Boundary Conditions
Phenomena which must be quantified Boundary layer
Blowing (if applicable decomposing materials) Velocity/temperature gradients Thickness
Shock interactions/effects Enhanced heating due to shock attachment (shock jetting)
Ionization (induces RF blackout, catalycity) Sodium/Potassium/air
Recovery conditions (enthalpy, temperature) Convective heat transfer (wall shear) Real/ideal gas effects
August 20, 2003 9
Candidate Material Technologies
Thermal properties (temperature/directional dependent) Thermal conductivity/specific heat Density (weight limits) Decomposition (char/pyrolysis) Ablation product species, ionization potential, catalytic
efficiency Mechanical properties (temperature dependent)
Strength (shear) Strain (motor growth/bending) Durability (impact, environmental extremes, )
Reusable or single use Cost/manufacturability/TRL
August 20, 2003 10
Analytic Model Development
Reusable technologies (NASA) Aeroheating + FEA Codes
Non-reusable (DoD) Aeroheating + Decomposition + FEA Codes Simplified approaches if applicable
Conduction models (unreliable if significant decomposition or ablation)
Simplified heat of ablation/Q* models (requires significant test data for specific environment of interest, limited use)
Complex charring material models Application to wider range of environments Requires characterization of complex phenomena Decomposition (char/pyrolysis/intumescence) Mechanical erosion/thermochemical ablation Properties dependent on temperature/char state/direction
August 20, 2003 11
Preliminary Flight Design Perform predictions with available material models
Aerothermal response (CMA) In-depth conduction (FEA) Note: Models may not exist or need development
Develop understanding of expected material phenomena/behavior for flight
Rank material performance (both TPS & Structure) Volume/weight constraints for required thickness
(substructure temperature limits) Application process (spray,trowel,mold,sheet) Cost Availability Technology readiness level (flight qualified?)
August 20, 2003 12
Ground test simulation of flight conditions Match aeroheating
Heat Flux (calorimeter) Shear Pressure Recovery conditions
Configuration (test rhombus) Instrumentation Screening tests
Facility availability Test cost
Ground Aerothermal Testing
Fwd ActuatorRear Actuator
Flow
Lasers
LoadcellStop Plate
Linear Guide
Test Plate
Fwd ActuatorRear Actuator
Flow
Lasers
LoadcellStop Plate
Linear Guide
Test Plate
August 20, 2003 13
Ground Aerothermal Testing Material performance characterization tests
Identify phenomena to characterize In depth conduction (thermal properties) Decomposition Erosion/thermochemical ablation Surface temperature Thermal expansion & durability
Select facility to characterize phenomena of interest
Thermal properties (thermal response) Decomposition Mechanical Erosion/Intumescence Ionization/plasma Heat flux/surface temperature response
Surface Recession /
IntumescenceChar
Pyrolysis
Virgin Layer
Substrate
Convection Net Radiation Blowing/Mass Injection
Boundary LayerFree Stream Flow
Chemical Diffusion/Reactions
Surface Shear
AblationConductionDecomposition
RX2390 TGA-DTA 50C/min 2/20/00
0
10
20
30
40
50
60
70
80
90
100
0 500 1000 1500 2000 2500 3000
Temperature (F)
Wei
gh
t (%
)
0
0.02
0.04
0.06
0.08
0.1
0.12
0.14
0.16
0.18
0.2
Wei
gh
t Lo
ss R
ate
(%/
F)
Weight %Weight Loss Rate (%/F)
0.09 Char
0.14 Pyrolysis
0.2 Virgin
0.43 TotalThickness
TGA/DSC
Intumescence &Decomposition
Area3Mean 718.0
Area3Mean 718.0
Area2Mean 726.0
Area2Mean 726.0
Area1Mean 778.9
Area1Mean 778.9
Area6Mean 842.9
Area6Mean 842.9
Area5Mean 854.1
Area5Mean 854.1
Area4Mean 885.9
Area4Mean 885.9
*>1,881F
*
August 20, 2003 14
Instrumentation Calorimeter matching test specimen Heat flux
Thin skin Plug calorimeter (copper,steel,water cooled) Heat flux gage (Schmidt-Boelter, Gardon, Null Point)
Flow field (pressure, temperature, boundary layer)
Embedded thermocouples Infrared surface temperature
(wavelength, emissivity) Erosion rate sensors Chemical species/injection
12
3
4 5 6 7 8 9
Area3Mean 718.0
Area3Mean 718.0
Area2Mean 726.0
Area2Mean 726.0
Area1Mean 778.9
Area1Mean 778.9
Area6Mean 842.9
Area6Mean 842.9
Area5Mean 854.1
Area5Mean 854.1
Area4Mean 885.9
Area4Mean 885.9
*>1,881F
*
August 20, 2003 15
Instrumentation
Transition Occurrence Altitude/velocity/AOA Surface roughness Critical for aerothermal
boundary condition definition
GageOnset
IndicatorMeasures
Q
Map Transition
Front
Measures Turbulent Level Q
CostTM
Bandwidth
Delta-T yes yes yes yes low low
Standard Calorimeter yes yes yes no low low
Ported Acoustic yes no could no high high
Unported Acoustic yes no could no med med
Base Pressure yes no no no low low
Accelerometer yes no partially no high high
Turbulent
Laminar
Table by Astrometrics
August 20, 2003 16Delta-T Gage To Determine Onset of Boundary Layer Transition Altitude
Minuteman Launch
Minuteman Re-entry Vehicle Quad Isothermal Plug Thermocouple
To Determine In-Depth Temperatures
ARAD sensor to measure the char virgin interface recession history
Transient Recession Rates
Non-Intrusive Embedded Thermocouples
Transition Occurrence
DOD Aerospace Off-the-ShelfHeat Shield Instrumentation
Instrumentation Charts provided by John Cassanto of Astrometrics (610) [email protected]
Instrumentation
August 20, 2003 17
InstrumentationTypical Flight System Instrumentation
August 20, 2003 18
Minimum Flight Instrumentation Requirements
Transition sensors (boundary conditions) Ablation sensors (heatshield response) Multi-element embedded plug TCs (TPS) Temperature probes (internal components) Backface structure temperature sensors Antenna window temperature sensors Dual range pressure transducers
August 20, 2003 19
Flight Design Verification and Validation
Utilize validated computational model developed from
ground test data to refine system design
Material requirements
Predicted thermal response
Predicted surface removal/intumescence
Substructure response (thermal/structural)
Flight instrumentation requirements
Telemetry capability (RF signal effects due to ablation
products)
August 20, 2003 20
Flight Design Verification and Validation
Data reduction
Final computational model refinement
Predictions for full range of possible flight
requirements
Ground tests cannot always fully match flight
Flight tests dont always impart worst case flight
August 20, 2003 21
Analysis and Design Methodologies
Summary of Methods Engineering Methods Complex Flow Field Analysis
Analytic Codes Aerothermal Boundary Conditions Material Response
August 20, 2003 22
Summary of Methods
Engineering Methods are relatively efficient and can provide a means of obtaining first order boundary conditions for simple configurations Panel methods Streamline tracing Equivalent running length Convective transfer expressions (Conical/flat plate/sphere/cylinder)
These methods can be supplemented with more complex flow field solutions to assess and correct pressure predictions and recalculate heat transfer coefficients
Engineering methods also provide efficient transient thermal response and boundary condition predictions with material shape change/ablation
Results can be integrated with FEA models for more complex 3-dimensional conduction effects
August 20, 2003 23
Computational fluid dynamics (CFD) codes may be used for complex and/or chemically reacting flows
Transient flow field predictions are generally not possible due to Length of solver run times Cost of discretizing domain for 3D geometry (grid) Inability to efficiently consider transient wall
temperature or ablation response Predictions are generally limited to a few points
in trajectory/time Used in a tabulated manner for verifying and
modifying engineering method predictions
Summary of Methods
August 20, 2003 24
Engineering Methods
Shape change Streamline tracing Surface pressure Shock shape Boundary layer flow
August 20, 2003 25
Shape Change
Geometry definition Freestream properties Inviscid flowfield
Surface pressure Shock shape
Boundary layer heating Material response and ablation Change in the geometry of the
vehicle Particle impact erosion Coupled shape change /
flight dynamics In-depth thermal response
Other factors Efficiency Robustness Accuracy
Inviscid flow
Boundary layers
Ablation
Shape change
Flight dynamics
Particle impact
August 20, 2003 26
Streamline Tracing
Axisymmetric analogy for 3D flowfield predictions Assumes no flow crossing a streamline Modeled as an axisymmetric solution Body radius replaced with the metric coefficient
Streamlines calculated using method of steepest descentbased on the Newtonian approximation
Newtonian flow model assumes that a stream of particles(air molecules) impinging on a surface retains its tangentialcomponent of momentum
August 20, 2003 27
Surface Pressure
h2* cospp CC =
R*
b*
Sonic point
II II III
Windward side (cos h < 0) (h is the angle between the wind vector and the outward surface normal)
Modified NewtonianPANT Correlations
Leeward sides (cos h >0)Newtonian Pressure:Small Disturbance TheorySeparation Correlations
0=pC
-
--=-
12
11
2)1(2
2
gg
ng
gM
MC p
Surface pressure calculatedalong streamline
h
q
August 20, 2003 28
Shock Shape
Thin shock layer approximation for bow shock solving the globalcontinuity and axial momentum equations. IntegrandsApproximated by linear distributions between the body surfaceand the shock.
p
u
vRu
rwr
d
y
Bow Shock
August 20, 2003 29
Boundary Layer FlowEntrainment relation - means of determining boundary conditions for the solutions of the momentum and energy equations. Edge state determined by lookup on pressure and entropy in a real-gas Mollier table. Pressure taken from inviscid pressure correlations and entropy is calculated by balancing mass entrained into boundary layer and mass crossing bow shock.
B o u n d a r y l a y e re d g e
r e u e s t r e a m l i n er u
yr w v w
August 20, 2003 30
Boundary Layer Flow
Transitional Boundary Layers:Transition from laminar to turbulent flow is modeled through theuse of the Persch intermittency factor. Once the transition criteriais met, MEIT calculates both laminar and turbulent solutions anduses this factor to determine the actual state of the boundary layer.The parameters Cf, Ch, H, F, and R are determined by the expression:
tPfPfP +-= l)1(
where f is the intermittency factor defined by:
)(Re1
,,2
lftf CCf
--=
q
a
wheretrftftr CC )(Re ,,
2, l-= qa
August 20, 2003 31
Boundary Layer Flow
Influence Coefficients:
Basic laws - represent simplest flowfield developing aboundary layer (incompressible flow, smooth wall, isothermal,
impervious, irrotational inviscid flow) Assumption - laws can be modified to account for nonideal
phenomena through use of influence coefficients These factors generally derived by comparing convective
transfer with the ideal flat-plate result for the sameboundary layer state.
The factors are derived for only one nonideal mechanismat a time
August 20, 2003 32
Boundary Layer Flow
===z
zyxyxyx tyfhxICC ,and,for,,0,,, l
where Cx,y,0 refers to the basic law for incompressible flow alongan impervious, isothermal flat plate
x indicates heat, h, or momentum, f, transfery indicates laminar, l, or turbulent, t, flowz indicates the nonideal effect being considered
b, accelerationB, blowingp, property and Mach number effectsr, roughness effectstr, transition proximity
It is assumed that the Stanton number, Ch, and the skin friction coefficient, Cf, can be written as:
Method utilized in ATAC3D - Influence Coefficients:
August 20, 2003 33
Complex Flow Field Analysis
Required due to geometry and/or flow physics Geometry complexity creates flow complexity
Shock-shock interactionsShock-boundary layer interactionWakesExpansion wavesFlow separation/recirculation/reattachmentChemistry/surface reactionsFlow injection
Engineering methods used to identify regions where CFD is required CFD utilized to provide refined boundary conditions for engineering methods
August 20, 2003 34
Analytic Codes for Boundary Conditions Engineering Methods
ATAC3D/MASCC Miniver/Lanmin IGHTS/RGHTS Bell Aeroheating Handbook
CFD Aerosoft - GASP CFDRC - Fastran Fluent Giants/LAURA (NASA - Axisymmetric) KIVA CFDL3 Overflow WIND
Coupled Design Codes IHAT (Miniver/ATAC3D/Overflow) Giants (TITAN, GIANTS, MARC-FEA)
Turbulent
Laminar
surface plot
FinenessRatio
InletPosition
MissileLength
Inlet Height,Width & Standoff
Fin Span, Sweep,Taper, Chord,
& Position
Diameter
Figure 2 : Air Launched Low Volume Ramjet configuration used as a validation test case.
Figure 7: Low-fidelity aerodynamic grid.
Figure 10: High Fidelity Aerodynamic Solution (Pressure coefficient at Mach 2.5, a =0)
August 20, 2003 35
Comparisons with Pegasus Flight Data
First Pegasus flights included instrumentation to measure interface temperatureson the wing, fin, and wing fillet.
Previous analytical models of Pegasus considered the fuselage, wing and fins asseparate entities. The current model includes the entire configuration.
Calculation modeled the first 82 seconds of the flight, up to first stage burnout 64 complete flowfield solutions were obtained to model the variation of the freestreamconditions during the flight.
CURRENT PEGASUS MODEL
HEAT FLUX CONTOURSAT THE END OF THEFLIGHT
August 20, 2003 36
CFD Solution with Adiabatic Wall
Trecovery
CFD Solution with Fixed Twall
Heat Flux
3D Thermal Analysis Stress Analysis
CFD, Thermal & Structural ModelingProcess for Complex Vehicles Shapes
Pressure
Temperature
Distribution
Other BCs
Time-dependent heat flux and Trecovery interpolated from CFD mesh to thermal mesh.
Time-dependent pressure distribution interpolated from CFD mesh to structural mesh.
Coupled CFD/Aerothermal Analysis
Technical POC: Rick Burnes Naval Air Warfare Center, [email protected]
August 20, 2003 37
Missile Aeroheating& Shape Change Code
I-DEAS TMGThermal Solver
TrajectoryGeometry/Mesh Geometry/Mesh/Materials/Initial BCs
Missile Skin Temps
hconvective , Trecovery
At each thermal solution time-step in a transient analysis: Missile skin temperatures and time-step passed to MASCC MASCC computes hconv and Trec using missile skin temperatures and trajectory for current time-step. Heat load on each surface element in thermal model at current time-step computed using: Qelement = (hconvective)(Aelement)(Trecovery - Telement)
Fortran User Subroutine Compiled & Linked to TMG
Coupled CFD/Aerothermal Analysis
August 20, 2003 38
Analytic Codes for Material Response
Charring Material Model CMA (Aerotherm)
CMA87: Charring Material/Ablation CMA92FLO: CMA87 + Pore Pressure
Numerous independently developed codes FIAT: 1-D NASA Charring Material Analysis TITAN: 2-D NASA Charring Material Analysis
Q* Model Simplified Heat of Ablation Model (modified Q* model)
Conduction Models
August 20, 2003 39
Charring Material Models Modeling Capabilities
In-depth decomposition Conduction Thermochemical ablation Fail temperature model for mechanical erosion Pyrolysis gas generation and injection/blowing Addition of intumescence for conduction effects
Application Generally required for decomposing materials to obtain
accurate in-depth thermal response Applicable to a wide range of environments once material
thermodynamic model is developed Limitations
Requires specific material property data such as kinetic decomposition constants, temperature dependent properties of char, pyrolysis, and virgin material, and heat of decomposition
August 20, 2003 40
Charring Material Thermal Response and Ablation and Model (CMA)
BoundaryLayer
Char
Pyrolysis Zone
Virgin Material
Backup Material
PyrolysisGases
ReactionProducts
Mechanical Removal-Melt Flow
-Particle Erosion
ConductiveFlux
RadiationFlux Out
RadiationFlux In
ConvectiveFlux
ChemicalSpecies
Diffusion
( ) ( ) 04** =-+----+- condstcwwseTwtcTiiwieTer qhmTFhmhmhZZMhhH www &&&& se
xhghz
TkA
zAz
Tpc
-+
=
qr
qqr 1
qqr
+
+zgh
Agm
zT
pcs&
&
Surface and In-depth Energy Balance
August 20, 2003 41
RX2390 TGA-DTA 50C/min 2/20/00
0
10
20
30
40
50
60
70
80
90
100
0 500 1000 1500 2000 2500 3000
Temperature (F)
Wei
gh
t (%
)
0
0.02
0.04
0.06
0.08
0.1
0.12
0.14
0.16
0.18
0.2
Wei
gh
t L
oss
Rat
e (%
/F
)
Weight %Weight Loss Rate (%/F)
( ) ( ) CA rrrr G-++G= 1Bi
i
i
i
i
o
rio
ai
x
i
RT
EB
j
r
rrr
qr
-
--=
exp
In Depth Decomposition Kinetics Model
August 20, 2003 42
CMA Intumescence Model Development
Substrate
Virgin
Pyrolysis
Char
SurfaceRecession
Intumescence
y
x
S
Boundary LayerFree Stream Flow
Surface Shear
Original Surface
Heated Surface
Net RadiationConvection
Ablation
Conduction
Decomposition
Blowing/Mass Injection
August 20, 2003 43
Intumescence Behavior
Pretest Thickness =0.3Char Depth =0.13Pyrolysis Depth =0.09Virgin Material =0.2
DepthPosttest Thickness =0.42Intumesced =120%
.420.420
.13 CHAR.13 CHAR
.20 VIRGIN.20 VIRGIN
.09 PYROLYSIS.09 PYROLYSIS
August 20, 2003 44
LHMEL Test Configuration
RTR device
Wind tunnel
exhaust tunnel
calibration block
sample holding fixture
RTR device
Wind tunnel
exhaust tunnel
calibration block
sample holding fixture
RTR device
Wind tunnel
exhaust tunnel
calibration block
sample holding fixture
LHMEL Facility
Test Hardware
Heatshield Samples Mounted Heatshield Sample
August 20, 2003 45
Posttest Samples
August 20, 2003 46
LHMEL Radiography Results
Thermocouples
Final SurfacePosition
Density Reference
Aluminum Substrate
Initial SurfacePosition
August 20, 2003 47
LHMEL Digitized RTR Results
CT Scan DataRX2390-S
1101704 75 S0932_03 1101704 1101704 1101705 1101705 1101706 11017061101704 75 S0932_03 1101704 1101704 1101705 1101705 1101706 1101706
0.0
0.5
1.0
1.5
2.0
2.5
3.0
-0.20 -0.10 0.00 0.10 0.20 0.30 0.40 0.50 0.60 0.70 0.80
Distance From Initial Flame Surface (in)
Den
sity
(g
m/c
c)
August 20, 2003 48
Comparison of Non-Intumescing and Intumescing Model Predictions
70
80
90
100
110
120
130
140
150
160
170
180
0 20 40 60 80 100 120 140 160 180 200TIME (SEC)
TE
MP
ER
AT
UR
E (
F)
TC2 ModelDATA TC2DATA TC3TC3-ARB KExisting Model
Modified Model Prediction
Test Data
Non-intumescing ModelInaccurate Slope
Intumescing ModelExceptional Agreement
August 20, 2003 49
NASA HGTF Low Shear Hypersonic TestsPhase 1 Cold Wall Conditions
Twall=70F
0
2
4
6
8
10
12
0 50 100 150 200 250 300 350
Time (sec)
Hea
t T
ran
sfer
Co
eff,
HO
(B
tu/f
t2-h
r-F
)C
old
Wal
l Hea
tin
g R
ate,
QC
W (
Btu
/ft2
-s)
0
500
1000
1500
2000
2500
3000
Rec
ove
ry T
emp
erat
ure
, TR
(F
)
HOQCWTR
Test #3 - RX2390 -2Position 4
0
100
200
300
400
500
600
700
800
900
1000
1100
0 20 40 60 80 100 120 140 160 180 200 220 240 260 280 300 320
Time (s)
Tem
pera
ture
(F)
TC1TC2TC3TC4TC5
Surface Temp
FlowTC3TC2
TC1 TC5TC4.344"
Density = 78.0 lb/ft3Delta T = 75.5 oFAverage Thickness = .344"Pre-Test Emissivity = .90Post-Test Emissivity = .92
Test #3 - RX2390 -2Position 4
0
100
200
300
400
500
600
700
800
900
1000
1100
0 20 40 60 80 100 120 140 160 180 200 220 240 260 280 300 320
Time (s)
Tem
pera
ture
(F)
TC1TC2TC3TC4TC5
Surface Temp
FlowTC3TC2
TC1 TC5TC4.344"
Density = 78.0 lb/ft3Delta T = 75.5 oFAverage Thickness = .344"Pre-Test Emissivity = .90Post-Test Emissivity = .92
0.13 CHAR
0.20 VIRGIN
0.09 PYROLYSIS
0.13 CHAR
0.20 VIRGIN
0.09 PYROLYSIS0.42 TotalThickness
0.13 CHAR
0.20 VIRGIN
0.09 PYROLYSIS
0.13 CHAR
0.20 VIRGIN
0.09 PYROLYSIS0.42 TotalThickness
Cold Wall Heating Rate
Measured Intumescenceand Decomposition
Test FacilityAnd
Configuration
Thermal Response
August 20, 2003 50
Model Thermal Predictions for HGTF Test
0
200
400
600
800
1000
0 50 100 150 200 250 300 350
Time (sec)
Tem
per
atu
re (
F)
Surface Predicted
0.05 Predicted
0.15 Predicted
0.25 Predicted
0.25" Data0.15" Data
0.05" Data
Surface Data
August 20, 2003 51
NAWC T-Range Ablation Tests
5.255
15
2.812
0.25
17-4 SS Tip
17-4 SS Substrate
Heatshield Sample17-4 SS Base
3.0
0.38
5.255
15
2.812
0.25
17-4 SS Tip
17-4 SS Substrate
Heatshield Sample17-4 SS Base
3.0
0.38
August 20, 2003 52
HHSTT High Shear Test Configuration
HeatshieldSample
17-4 StainlessSteel Tip
17-4 StainlessSteel Substrate
17-4 StainlessSteel Base
15 Conical Sample
2.12
HeatshieldSample
17-4 StainlessSteel Tip
17-4 StainlessSteel Substrate
17-4 StainlessSteel Base
15 Conical Sample
2.12
0
100
200
300
400
500
600
0 1 2 3 4 5 6 7 8 9
Time (sec)
Col
d W
all H
eat F
lux,
q (B
tu/ft
2-se
c)
0
200
400
600
800
1000
1200
Inte
grat
ed C
old
Wal
l Hea
t Flu
x, Q
(Btu
/ft2)
q
Q
LX=3"Alpha=15Tcw=80F
August 20, 2003 53
Coupled Intumescence and MechanicalErosion Modeling
-0.05
0
0.05
0.1
0.15
0.2
0 10 20 30 40 50Time (sec)
Su
rf P
osi
tio
n (
in)
T-Range, Tc=1600F
T-Range, Tc=1000F
LHMEL 75 W/cm2
August 20, 2003 54
Q* Models Modeling Capabilities
Ablation measured to virgin/pyrolysis interface(ablation temperature, heat of ablation)
Conduction Blowing for sublimation process Simplification of required input parameters
Application Generally applicable to materials/environments with high
mechanical erosion/ablation Not appropriate for highly decomposing/intumescing materials
Limitations Requires ablation performance test data for heating/shear
environment of interest Can provide misleading results if extrapolated to other
environments Does not necessarily provide realistic surface temperature response
or physical surface removal prediction (simplified heat of ablation model)
August 20, 2003 55
Heat of Ablation Model Development
Virgin MaterialVirgin
Material
SubstrateSubstrate
AblatedDepth
AblatedDepth
Original SurfaceOriginal Surface
Radiative Heat
Radiative HeatConvective
HeatConvective
Heat
BoundaryLayer
BoundaryLayer
SSAblationAblationConductionConduction
Energy Balance
Ablation Surface Boundary Condition
=
xT
kxt
TC pr
( ) HstqxT
k acsx ==
-
= r
)(;)(
ablationofperiodduringa
dtTTh
aq
Hatotvirgin
ar
totvirgin *
-=
*=
rr
August 20, 2003 56
Typical Heat of Ablation Model Development
Heat of Ablation, Ha (Btu/lbm)
To
tal A
bla
tio
n D
epth
, ato
t (in
)
0
0.05
0.1
0.15
0.2
0.25
0.3
1000 1200 1400 1600 1800 2000 2200 2400 2600 2800 3000
0.098" Total Ablation
Ta=800 HA=1310
Ta=500 HA=2040
Ta=300 HA=2500
Atot data
Legend
August 20, 2003 57
Ha Ta for Phase 1
T backside
Tb Measured
Legend
Typical Heat of Ablation Model Development
Phase 1 RX2390
0
200
400
600
800
1000
1200
1400
1600
1800
2000
2200
2400
2600
2800
3000
100 200 300 400 500 600 700 800 900
Ablation Temperature (F)
Hea
t of A
blat
ion
(Btu
/lbm
)
100
150
200
250
300
350
400
450
500
550
600
Ste
el B
acks
ide
Tem
pera
ture
(F)
Ha=2720 Btu/lbmTm=192F
Ha
Ta
August 20, 2003 58
Time (sec)
Typical Heat of Ablation Model Development
Tb Simplified HOA
Tb Measured
Legend
Phase 1 RX2390
70
90
110
120
140
150
170
0 50 100 150 200 250 300 350
Bac
ksid
e T
emp
erat
ure
(F
)
q totalheat in
q internalradiation
Thermocoupletemperature
Atot
Virgin materialq totalheat in
q internalradiation
Thermocoupletemperature
Atot
Virgin material
Inaccurate slope ofthermal responsefor simplified heatof ablation model
80
100
130
160
Time (sec)
August 20, 2003 59
Conduction Models
Modeling Capabilities Conduction Modified thermal properties to account for other thermal
response phenomena (density change) Application
Generally applicable to materials/environments with no decomposition or surface removal/ablation
Common approach for finite element analysis until a coupled ablation capability is developed
Limitations Does not allow surface removal Does not generally account for density changes that effect
conduction Can provide misleading results if used for
ablating/decomposing materials
August 20, 2003 60
Examples of Aerothermal Analysis and Design
Missile Electronics Unit
InertialMeasurement Unit
Control ActuatorFin Assembly
Airframe FEThermal Model
Missile Electronics UnitCard Cage
Missile Electronics Unit
Guidance Electronics
Control Actuator Fin Assembly
Inertial Measurement Unit
VirtualPrototyping
Analysis/Simulation
DetailedDesign
Simulation BasedDesign
August 20, 2003 61
Optical WindowCandidate materials analyzed
ALONSapphireDynasil
Bevel angles analyzed45 and 60 degree
FrameSteel
ReflectorAluminum
Grafoil- Used at interface between Window and Frame to provide a seal and minimize stress. Also used between Window and Reflector to distribute seal pressure loads.
Detector HousingAluminum
2366
0
500
1000
1500
2000
2500
3000
0.0 0.5 1.0 1.5 2.0 2.5 3.0 3.5 4.0
Time (sec)
Rec
ove
ry T
emp
erat
ure
(d
eg-F
)
Baseline
Test Flight
T-Range
Window
Frame
Grafoil
Detector Housing
1 Distance
Test ArticleT-Range 9 Nozzle T-Range Sting
1 Distance
Test ArticleT-Range 9 Nozzle T-Range Sting
12
34 5 6 7 8 9
Examples of Aerothermal Analysis and Design
System Requirements Flight Environments
AerothermalAnalysis
Material Selection
Ground Test& Evaluation
Flight Test Verification& Validation
August 20, 2003 62
Summary
This tutorial has provided a brief overview of the aerothermal analysis and design process
Various components of this process have been discussed
A limited survey of existing methodologies and corresponding codes have been provided
August 20, 2003 63
Summary
The goal of this tutorial was to define a general process in which aerothermal analysis and design should be conducted.
It is imperative for the designer to consider a wide range of issues when conducting aerothermal analysis and design.
Flexibility should be maintained during the process to ensure critical issues are adequately resolved prior to system final design.
Engineering methods represent an efficient approach to design. However, more complex approaches should be utilized to supplement these engineering methods when deemed necessary.