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

*Gerald.russell@rdec.redstone.army.mil

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) 280-0869AstrometricsJMC@aol.com

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, BurnesRR@navair.navy.mil

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.