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Hypersonic Vehicle Systems Integration Vehicle Aerothermodynamic Analysis and Thermal Management Des

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Ablator. Insulation. Structure. AEROHEATING. Radiative Heat ... Ablator TPS used on earth/planetary re-entry vehicles or ballistic/hypersonic weapon systems ... – PowerPoint PPT presentation

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Title: Hypersonic Vehicle Systems Integration Vehicle Aerothermodynamic Analysis and Thermal Management Des


1
Hypersonic Vehicle Systems IntegrationVehicle
Aerothermodynamic Analysis and Thermal Management
Design
  • 12 September, 2007
  • Dr. Kevin G. Bowcutt
  • Senior Technical Fellow
  • Chief Scientist of Hypersonics
  • Boeing Phantom Works

2
Different Flight Regimes Influence Nature of
Aerodynamic Heating and Analysis Requirements
Pressure and viscous effects important Transition
laminar-turbulent Low density effects Surface
radiation Strong reacting flow effects
Speed Regimes
Flow Properties
Limit of continuum regime
100
Subsonic (M lt 0.8)
Ideal Gas
Slip-flow regime
Transonic (0.8 lt M lt 1.2)
Weak Shock Waves Ideal Gas
Ascent and Reentry Vehicle (ARV)
Reentry Vehicle (RV)
h (km)
Aeroassisted Orbital Transfer Vehicle (AOTV)
Pressure effects dominate Strong reacting flow
effects Surface radiation Low density effects
Supersonic (1.2 lt M lt 5.0)
Shock Waves Calorically Imperfect Gas
Predominantly laminar flow
50
Ionization Radiation Strong reacting flow
effects Low density effects
Strong reacting flow effects
Hypersonic (M gt 5.0)
Thin, Hot Shock Layers Thermally Imperfect Gas /
Chemically Reacting
Laminar-turbulent transition
Stage separation of Two-Stage-to-Orbit (TSTO)
system
  • Flow Regimes (Knudsen Number, Kn)
  • Continuum Flow (ÖRe/Mgt100 and Knlt0.01)
  • Slip Flow (1ltÖRe/Mlt100 and 0.01ltKnlt0.1)
  • Transitional Flow (0.1ltÖRe/Mlt1 and 1ltKnlt10)
  • Free Molecular Flow (ÖRe/Mlt0.1 and Kngt10)

Viscous effects large Laminar-turbulent
transition Weak real-gas effects
Cruise Vehicle (CV)
0
0
1
2
3
4
5
6
7
8
9
10
U? (km / s)
Adapted from Bertin 2
  • Most severe heating problem occurs in continuum
    regime
  • Imperfect gas effects become important at Mach gt
    3
  • Significant O2 and N2 dissociation occur in
    high-temperature flow
  • Thermochemistry model required for aero-heating
    analysis
  • Rarefied gas effects occur in high altitude
    flight regime

3
Thermal Protection and Management at Hypersonic
Conditions is Challenging
  • Typical surface temperatures illustrated by 20
    wedge 10-feet from leading edge (? 0.8)
  • Flow specific energy increases as V2
  • Aerodynamic heating (energy flux) increases as V3
  • Boundary layer state an important driver of heat
    flux wall temperature

4
Typical Aero-Heating Environmentof Hypersonic
Vehicles
Space Access Vehicles
Atmospheric Cruise Vehicles
Nose Cap 1200 Btu/ft2-s
Nose Cap 700 Btu/ft2-s
Leading Edge 900 Btu/ft2-s
Leading Edge 500 Btu/ft2-s
Inlet Ramp 100 Btu/ft2-s
Inlet Ramp 50 Btu/ft2-s
Cowl Lip 3500 Btu/ft2-s
Engine Internal 1200 Btu/ft2-s
Engine Internal 1500 Btu/ft2-s
Cowl Lip 2000 Btu/ft2-s
External Nozzle 180 Btu/ft2-s
Nozzle 60 Btu/ft2-s
w/o Shock-Shock Interaction
5
Surface Aerodynamic Heating
  • A correlation exists between velocity and
    temperature gradients in boundary layers called
    Reynolds Analogy that permits estimating
    convective heat flux from frictional shear stress
  • Radiation equilibrium wall temperature can then
    be found by equating convective and radiative
    heat fluxes

6
Friction Drag Estimation for Predicting Heat Flux
  • Use flat plate theoretical formulas with
    empirical reference-temperature corrections for
    high speeds

Laminar
Turbulent
7
Stagnation Point Aerodynamic Heating
  • Self-similar laminar boundary solutions produce
    equations for stagnation point convective heat
    flux on cylinders and spheres
  • Since stagnation point heat flux is inversely
    proportional to the square-root of leading edge
    radius, it is common practice to compute the heat
    flux on a unit cylinder or sphere along the
    flight trajectory, allowing heat flux to be
    readily scaled for different radii
  • Stagnation point wall temperature can be found by
    equating convective and radiative heat fluxes,
    the same as for surface heating

8
Shock Interactions Substantially Increase Local
Heating
  • Localized aerodynamic heating created by shock
    interactions can result in TPS material failure
    if heating magnitude not adequately predicted

Corner Shock Interaction (End View)
Swept-Fin Mounted On Plate
Separation Reattachment
Shock / Boundary-Layer Interaction (Flow
Deflection)
Viscous Interaction
Corner Flow
Shock/ Boundary-Layer Interaction (Flow
Deflection)
Wing Leading-Edge Shock / Bow Shock Interaction
Jet Interaction
Type of Boundary Layer
Corner Flow With Shock Impingement
Bow Shock / Cowl Shock Interaction
Strut Interactions
Shock / Shock Interaction (Shock on Lip)
Inlet Spillage Boundary Layer Ingestion
Ramp Shock Interactions
From Bertin 2
9
Regions of Strong Viscous Interactions
  • Shock / shock and shock / boundary layer
    interaction can magnify heating rates dramatically

Glancing Interaction On Roof and Floor Caused By
Fuel Injection Strut
Shock / Shock Interaction
Numerical Experimental
M? gt gt 1
Laminar Viscous Interaction
10
10
8
8
Glancing Interaction On Side Walls
Oblique Shock Boundary Layer Interactions
and Turbulent Viscous Interaction
Transition
6
6
4
4
Bow Shock-Wave
M2 lt 1
2
2
Jet Bow-Shock
Slip Line
M4 gt 1
2
Rb
0
0
E
A
C
-2
-2
1
Body Surface
7
4
-100
-50
0
100
50
-100
-50
0
100
50
8
M1
6
F
? ()
? ()
B
D
Impinging Shock-Wave
Jet
(a) Surface pressure
(b) Heat transfer
3
5
(?sw)13
Surface-pressure and heat-transfer-rate
distributions for a Type IV shock / shock
interaction, M1 8.03, Re / ft 1.55 x 106
M5 lt 1
?13
Shock Generator
Type IV shock / shock interaction pattern
From Bertin 2
10
Shock-Shock Interactions on Leading Edges Can
Severely Magnify Aero-Heating
Van Wie, David M., Purdue University Short
Course, NASA Marshall Space Flight Center, Feb.
7-11, 2000
11
Typical Thermal Protection and Thermal Management
System Concepts
Passive
High Temperature Heat Sink Panel Example
Metallic TPS, TMC
Flexible TPS Blanket Example AFRSI, AETB
Rigid Ceramic Tiles Example FRCI, AETB
Temperatures Up to 1200F Application Leeward
Surface
Temperatures Up to 1500F Application Leeward
Surface
Temperatures Up to 2800F Application Windward
Surface
Carbon Ceramic Composite Example ACC, C-SiC
Ultra-High Temperature Ceramic Example Zirconium
Diboride
High Temperature Dielectric Composites Example
Dynasil, 3DQ DI-200
AEROHEATING
Radiative Heat Loss
C-SiC Skin Insulation
Structure
FUEL HEAT SINK
Temperatures Up to 3000F Application Windward
Surface, Control Surfaces, Blunt Nose/LE
Temperatures Up to 3700F Application Sharp
Nose/LE
Temperatures Up to 4400F Application Antenna
Windows
12
Typical Thermal Protection and Thermal Management
System Concepts
Active TPS
Semi-Passive TPS
Heat Pipe
Filming Cooling
Transpiration Cooling
AEROHEATING
Air Flow
Coolant Flow
Temperatures N/A Application Nose/LE, Engine
Parts
Temperatures N/A Application Nose/LE, Engine
Parts
Temperatures N/A Application Nose/LE, Engine
Parts
Ablator Example 3-D C-C
Actively Cooled Panel
  • Engine Heat Loads gt 80 of Total
  • Cooling Capacity gt Heat Loads

AEROHEATING
Radiative Heat Loss
Cooling Capability
Cooling Capability or Heat Load (kBtu/Sec)
Flow
M 0.8 Cruise
Ablator Insulation Structure
Turn
Cruise
Descent
Total Heat Load
Temperatures N/A Application Usually Not
For Space Plane
Temperatures N/A Application Nose/LE, Engine
Parts
7000
6000
5000
4000
3000
2000
1000
0
Time (Sec)
13
Thermal Protection System Requirements
TPS
  • Low thermal conductance allows complex TPS
    concept to be represented by a 1-D thermal
    network (i.e., small in-plane heat transfer)
  • External boundary condition dependent on class of
    TPS (ablative or non-ablative)
  • Internal boundary condition dependent on location
    or application
  • Limit structure temperature to material allowable
    (all locations)
  • Limit heat flux to avoid excessive fuel boil-off
    (cryogenic fuel tank)
  • Limit heat flux to internal bays (crew station,
    payload or equipment bay)
  • Thermal analysis conducted across entire mission
    profile to verify thermal constraints are met, or
    to re-size TPS to meet constraints

Courtesy of Kei Lau, Boeing
14
First-Order Method for Predicting TPS Inner
Surface Temperature As a Function of Time
  • Requires solution of one-dimensional version of
    Fouriers law of heat diffusion for a single
    material with uniform material properties

w
TPS or Structure
Z
15
Modern Reusable TPS and TPS Modeling Examples
Alumina Enhanced Thermal Barrier (AETB) Tiles
with TUFI Coating
Titanium Multi-wall Thermal Protection System
Tailorable Advanced Blanket Insulation (TABI)
Superalloy Honeycomb Metallic TPS
  • Analysis of 1-D thermal network determines TPS
    thickness (t) required

Courtesy of Kei Lau, Boeing
16
Ablative TPS Has More Complex Thermal Response
  • Ablator TPS used on earth/planetary re-entry
    vehicles or ballistic/hypersonic weapon systems
  • Thermo-chemical reaction occurs under high heat
    flux
  • Material composition changes
  • Moving boundaries between layers
  • Mechanical erosion of outer layer possible
  • Current state-of-the-art analysis methods
    deficient

Courtesy of Kei Lau, Boeing
17
1968 Apollo TPS Design Example
  • AVCO 5026-39G
  • Epoxy-novalac with Quartz fiber Phenolic
    Micro-balloon
  • Gunned into honeycomb matrix
  • HT-434 Phenolic adhesive bonded to SS-steel shell

NASA TN D-7564, JSC Apollo Experience
Report-TPS, J E Pavlosky and L G St. Leger, 1973
  • Success attributed to conservative design
    philosophy and rigorous testing
  • Monolithic heat shield required lengthy
    development and special manufacturing
  • Extensive arcjet tests for material
    characterization
  • Ground thermal-vacuum and flight tests used for
    system qualification
  • Separate flights for peak heat rate and maximum
    total heat load
  • High uncertainty in aerothermal environment in
    regions of penetrations, cavities and
    protuberances resulted in design conservatism

Courtesy of Kei Lau, Boeing
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