2. American Institute of Aeronautics and Astronautics
2
L Length, inches or feet
M, Ma, Mach Mach Number
Me Boundary Edge Mach Number
ML Boundary Edge Mach Number
pe Pressure at the Boundary Layer Edge, psia or psfa
pt, pt1 Total or Settling Pressure, psia or psfa
pt3 ET Nose Spike Pressure, psia or psfa
p∞, p Free Stream Static Pressure, psia or psfa
Prw Prandtl Number at the Wall
q∞ Free Stream Dynamic Pressure, psi or psf
r Recovery Factor
R Radius, ft
gR Gas Constant per Unit Mass
Re Reynolds Number
Reθ Momentum Thickness Reynolds Number
St Stanton Number
Stref, St (ref) Stagnation Point Stanton Number
St∞ Free Stream Stanton Number
Taw Adiabatic Wall Temperature, °R
To, TO, To Total Temperature, °R
Tt Total Temperature, °R
Tw, Tw Model Wall Temperature, °R
T∞, T Free Stream Static Temperature, °R
v∞, v Free Stream Velocity, ft/sec
x/L Longitudinal Position as a Fraction of the Reference Length
α, Alpha Angle of Attack, degrees
β, Beta Angle of Sideslip, degrees
γ Ratio of Specific Heats, cp/cv
µ∞, Mu Free Stream Dynamic Viscosity, slugs/ft-sec
µe Dynamic Viscosity at the Boundary Layer Edge, slugs/ft-sec
µw Dynamic Viscosity at the Wall, slugs/ft-sec
ν Kinematic Viscosity, slugs/ft-sec
θ Boundary Layer Momentum Thickness
ρ∞, Rho Free Stream Density, slugs/ft3
ρe Density at the Boundary Layer Edge, slugs/ft3
Subscripts
aw Adiabatic Wall Condition
e Boundary Layer Edge Condition
i Subscript Denoting Local Conditions
inf Free-Stream Condition
L Boundary Layer Edge Condition
o Total Condition
ref Stagnation or Reference Condition
w Wall Condition
∞ Free Stream Condition
I. Introduction
The Hypersonic International Flight Research and Experimentation (HIFiRE) program is a joint effort between
the United States Air Force Research Laboratory (AFRL) and the Australian Defence Science and Technology
Organisation (DSTO)1
. Boeing is an industrial partner in the HIFiRE program, having joint responsibility with
DSTO for three flight experiment payloads (numbers four, six, and eight). The overall objective of the HIFiRE
program is to gather fundamental scientific and engineering data on the physics and technology that is critical to
DownloadedbyKeiYunLauonNovember7,2016|http://arc.aiaa.org|DOI:10.2514/6.2012-5842
3. American Institute of Aeronautics and Astronautics
3
future operational hypersonic vehicles. This is accomplished by conducting several flight experiments using a
relatively low cost sounding rocket based flight test technique advanced by the researchers from DSTO.
HIFiRE 4 is an unpowered, waveriding, lifting body, flight mechanics experiment. Boeing was tasked to develop
an aerodynamic design database and the preliminary level design of the vehicle2
. Boeing generated a database of
7200 flight conditions using 1000 Euler CFD runs and Design-of-Experiments software to predict the remaining
6200. Boeing incorporated a complete CFD lateral-directional database within its simulation environment and
performed stability and control verification simulations. Boeing defined avionics and subsystem configurations in
conjunction with MATLAB Simulink models suitable for autocoding of flight control and guidance of an enhanced
HIFiRE 4 mission within the atmosphere. Boeing performed simulation trades and analysis to define flight control
hardware requirements for the free flight experiment. Boeing generated Hardware in the Loop (HIL) simulation
models of the free-flight experiment. Boeing evaluated via the results of existing thermal analysis the need for
acreage thermal protection on the free-flight configuration. Boeing completed a trade study of a blunt aluminum
leading edge versus a sharp copper leading edge and concluded that copper was sufficient and superior. Boeing then
proceeded to design and fabricate two shipsets of copper leading edges for the HIFiRE 4 flight experiment. Other
details of the leading edge design will be discussed in later section.
The subject of this paper is the design approach of this low cost airframe for flight test. The HIFiRE-4 vehicle
flight profile presented additional challenges in airframe design that are different from a ballistic entry. The HIFiRE-
4 airframe is constructed of conventional metallic materials including aluminum and copper. It is highly desirable to
have a fully three dimensional thermal and structural analyses to verify the integrity of the airframe design,
particularly the sharp leading edges which were subject to high aerodynamic heating. This paper describes the
process which was accomplished with limited resources.
The Boeing Company Phantom Works organization has developed several hypersonic designs since the early
1980's including the on-going X-51A SED3
. In 2005, NASA and its ATK/Boeing contractors completed the
successful Hyper-X (X-43A) flight test program4
. The aerodynamic environment can be characterized as extremely
harsh and the vehicles often function near the edge of system capability. The design of hypersonic systems requires
major advances in physical understanding and predictive capability in a number of areas, including air-breathing
propulsion, thermal protection and management, control authority for maneuver and precision targeting, and
boundary layer transition prediction and control. Figure 1 shows some recent hypersonic Programs at Boeing.
DownloadedbyKeiYunLauonNovember7,2016|http://arc.aiaa.org|DOI:10.2514/6.2012-5842
4. American Institute of Aeronautics and Astronautics
4
Figure 1, Some Recent Boeing Hypersonic Programs
II. Airframe Design Challenges and Conceptual Trade Study
A more detailed description of the HIFiRE-4 flight experiment objective and the vehicle configuration is in
Reference 2. The primary objective of HIFiRE 4 is to pull out from a near vertical atmospheric re-entry through an
angle of 25 degrees. The secondary objective is to then bring the vehicle to the ground via a portion of horizontal
flight. This flight will gather flight data on the aerodynamics, stability, and control of an advanced waverider
configuration performing a pull-up maneuver from a near vertical atmospheric entry flight path angle to horizontal
flight. The configuration was developed around an osculating cone waverider, with a cylindrical fuselage added to
the upper surface for housing subsystems and the outboard portions of the waverider wing truncated to fit within a
payload shroud without having to scale the vehicle down to a size too small to allow all flight test objectives to be
met. Vertical stabilizers were added to the truncated wing tips for directional stability and to maintain waverider
pressure levels on the wing undersurface. The resulting HIFiRE 4 configuration is shown in Figure 2.
DownloadedbyKeiYunLauonNovember7,2016|http://arc.aiaa.org|DOI:10.2514/6.2012-5842
5. American Institute of Aeronautics and Astronautics
5
Figure 2, HIFiRE-4 Configuration and Component Layouts
The HIFiRE-4 trajectory begins with a rocket boost out of the atmosphere, followed by a 25-degree angle of
attack pull-up maneuver to attain horizontal flight after atmospheric entry. The pull-up maneuver decelerates the
vehicle from Mach 8 to Mach 4 in about 30 seconds.
Earlier HIFiRE vehicles were constructed of relatively low temperature conventional materials such as
aluminum without external thermal protection system. This approach was demonstrated successfully on HIFiRE-15
.
The HIFiRE-1 is an axi-symmetric configuration that flew a straightly ballistic trajectory. Can this low cost airframe
approach be applied to a non-axisymmetric configuration that flies a pull-out maneuver for longer duration? What
changes will need to be made to accommodate the high aerodynamic heating of a sharp leading edge? Early in the
program, a preliminary thermal analysis was conducted to assess the adequancy of conventional materials for this
mission. The candidate materials were aluminum for the body and copper for the leading edges. The approach is
similar to, but modified from a transitional thermal protection system analysis such as described in Reference 5. The
objective of HIFiRE-4 design is to use the thermal capacitance of the airframe material as thermal protection. If
successful, the HIFiRE-4 airframe construction cost will be a small fraction of conventional hypersonic flight
vehicle, such as X-51A, that required carbon-carbon leading edge and complex external TPS design.
A. Engineering Aeroheating Method
The aeroheating analysis was conducted using the MINIVER engineering aeroheating code. This code is based
on impact theories and Reynolds analogy extrapolated skin friction point-to-point correlations7
. Different flow
models for various sections of the flight vehicle is required to mimic the flow field phenomena observed in
experimental data and 3-D CFD results of this class of configuration. Extensive calibration has been done to
establish the accuracy of the flow field models and heating methods. Corrections to the shock, pressure and heating
methods have been derived from the NASP (National AeroSpace Plane) ground test database8
. Behind-shock and
local flow properties based on oblique shock or sharp cone entropy conditions are determined based on selected
shock shape and pressure options. The accuracy of the aerothermal environment predictions using an engineering
tool is influenced by the experience of the analyst9
. The boundary layer transition prediction method adapted for the
aerothermal analysis is based on methodology and analytical studies developed during the NASP program. Details
of the development and background information can be found in Reference 10. A process has been developed to
automatically extract geometry data required in the aeroheating analysis from any vehicle model using the CATIA-5
CAD design tool. Flight trajectory data including altitude, Mach number and angle of attack were also required for
the analysis.
The two HIFiRE-4 flyers will be launched together on a VSB30 sounding rocket. The mission is shown in
Figure 3. At 120 km the two flyers are released (point 8). They coast to 292 km maneuvering with the exo-
atmospheric control system (point 9). Once they descend to 100 km the aero-control system takes over (point 11).
At 88 km the pull-up maneuver begins (point 12) and by 29 km the DSTO experiment ends. The DSTO flyer then
DownloadedbyKeiYunLauonNovember7,2016|http://arc.aiaa.org|DOI:10.2514/6.2012-5842
6. American Institute of Aeronautics and Astronautics
6
falls to an impact in the Woomera Test Range. At 29 km the Boeing secondary experiment begins (point 13). At
that point the flyer transitions from the DSTO flight control algorithm to the Boeing algorithm. The vehicle
executes a pull-up to horizontal and then impacts Woomera more than 300 km downrange. The re-entry aeroheating
analysis started from the Mach 8, 88 kilometer point to payload landing. There is very little aeroheating effect after
the pull-up (Point 15). Figure 4 shows the cold wall (0°F) aeroheating distribution at 27 seconds after entry where
the peak heating occurs. Neither CFD solutions nor aerothermal wind tunnel data were used due to limited budget
and quick turn around schedule. It is the authors opinion that the MINIVER only approach is deemed adequate for
the short flight duration and an all metallic airframe design based on his experiences.
Figure 3, HIFiRE-4 Notional Mission Profile
DownloadedbyKeiYunLauonNovember7,2016|http://arc.aiaa.org|DOI:10.2514/6.2012-5842
7. American Institute of Aeronautics and Astronautics
7
Figure 4, Peak Aeroheating Distribution at 27 Seconds after Entry
B. Quick Turn-around Thermal Modeling
In a conventional hypersonic airframe design, the first step is the selection of external thermal protection (TPS)
material and its sizing. The analysis is usually performed with 1-D thermal modeling with accurate material models,
particularly for ablative TPS. The HIFiRE-4 design objective is to use the conventional materials airframe as the
heat sink. The high thermal conductivity of the airframe material made it necessary to account for the lateral
conduction effect because a 1-D modeling approach will be too conservative.
At this point of the design, the HIFiRE-4 configuration is frozen with internal details partially defined. A pseudo
3-D thermal model was created using a 2-D CATIA geometry file. The model has copper leading edges and
aluminum panels, but no bulkhead or other internal details. Internal walls between the body and wing were modeled.
The total thermal masses were simulated using smeared panel thicknesses. Figure 5 shows a cutout aft section of
the model.
Figure 5, Preliminary Pseudo 3-D Thermal Model
Internal wall between body and wing modeled
Smeared wing panel thickness
DownloadedbyKeiYunLauonNovember7,2016|http://arc.aiaa.org|DOI:10.2514/6.2012-5842
8. American Institute of Aeronautics and Astronautics
8
Multiple analyses were traded with different body panel thicknesses. It was determined that a 5.6 m.m. body and
wing panel thickness would provide sufficient protection. The vertical fin panel thickness was minimized to help
control C.G. A 3 m.m. thick panel was deemed adequate for the fin. The effectiveness of thin thermal protection
coating using Sparesyl and phenolic cork were investigated. But the simplicity of bare aluminum panel is preferred
and the required thickness can be accommodated. Figure 6 Shows the temperature distribution with required panel
thickness.
Figure 6, Panel Temperature at 34 Seconds After Entry
Figure 7, Leading Edge Fine Grid Thermal Models
The leading edge of HIFiRE-4 is subjected to the highest aeroheating with a peak value (cold wall) of 380
Btu/ft2
-sec. A fine grid 2-D thermal analysis was performed to answer the questions: (1) Can the copper leading
°F
Cut #1
Cut #2
Copper in Green
Fine grid at nosetip
Cut #1
Cut #2
DownloadedbyKeiYunLauonNovember7,2016|http://arc.aiaa.org|DOI:10.2514/6.2012-5842
9. American Institute of Aeronautics and Astronautics
9
edge meet the mission requirement? (2) Will the leading edge temperature drop off rapidly enough from stagnation
to achieve an allowable aluminum temperature at the attachment points? Figure 7 shows the locations of the model
cut and the grid distribution. The highest stagnation aeroheating occurred at the centerline (cut #1) location, but it
has a thickness include angle to offset the effect. The off-center (cut #2) aeroheating is lower, but the leading edge
include angle is much smaller. Figure 8 shows the time history and peak temperature distribution result at cut #2.
The peak stagnation temperature at cut #1 is 600°F lower. It is noted that three dimensional conduction is not
rigoriously represented in the 2-D thermal models. As a result, the off centerline wing leading edge temperature is
likely over-predicted and the centerline fuselage leading edge temperature is optimistic. It is concluded that the
copper leading edge in current configuration can meet the mission requirements.
Figure 8, Copper Leading Edge Temperatures
These preliminary analyses allowed the Boeing design team to proceed to detail design with confidence.
III. Final Thermal Modeling of HIFiRE-4
The HIFiRE-4 structural concept is a one-piece wing machined from 6061-T6 Aluminum plate with thinner
removable covers fastened on as needed. The leading edges are segmented C11000 Copper. The simplicity of the
concept is similar to wind tunnel model practices. It affords the flyers to have a relatively thin wing and it reduces
manufacturing costs. The extra weight incurred by not having a built up wing doesn’t exceed the booster allowances
and doesn’t hurt the performance of the experiment either. Using the guidelines developed in Section II, the
aluminum bulk serves the additional purpose of being a heat sink to handle the thermal loads. Figure 9 shows the
HIFiRE-4 structural components and the internal subsystem placement.
A fully three dimensional thermal model was created directly from the CATIA-5 CAD drawing files with all
internal details into I-DEAS (Integrated Design and Engineering Analysis Software by Siemens). The surface
geometry was meshed for radiation calculation and external aeroheating mapping. The IDEAS/TMG pre-processing
tool provides accurate calculation of thermal mass and conductor couplings for both thermal and radiation models.
The pre-processing results were exported to TRASYS for radiation analysis and to SINDA for thermal analysis. The
aeroheating approach is the same as described in Section II. The aeroheating results were then mapped to the surface
of a 3-D thermal model for thermal analysis with SINDA. A mapping tool automatically maps aero-heating results
generated from the aero-heating module to the 3-D thermal model. The mapping tool reduces the overall design
cycle time and most importantly reduces analysis errors. The overall aerothermal and thermal analysis approach is
0
200
400
600
800
1000
1200
1400
1600
0 5 10 15 20 25 30 35
Time (Seconds)
Temperature(°F)
Nose Tip
FS 3.3"
°F
DownloadedbyKeiYunLauonNovember7,2016|http://arc.aiaa.org|DOI:10.2514/6.2012-5842
10. American Institute of Aeronautics and Astronautics
10
the same as decribed in Reference 8 and illustrated in Figure 10 in graphic form. Figure 11 shows an example of the
surface heat flux mapping results.
The HIFiRE-4 vehicle reaches peak temperatures at 35 seconds after re-entry, lagging the peaking aeroheating by
7 seconds. The surface panel temperatures and internal temperature distribution at 35 seconds are shown in Figure
12 and Figure 13, respectively.
Figure 9, HIFiRE-4 Structural Component Layout and Subsystem Placements
Figure 10, HIFiRE-4 Detail Thermal Analysis Flow Diagram
FEM Solid
Modeling
Convert FEM into
Thermal Analysis
Model
External
Surface Grid
Generation
(CATIA)
Trajectories
External Surface
Geometry and
Streamlines
Convective Heat
Transfer
Baseline
Configuration
(CAD Model)
Transient Finite
Difference
Thermal Analysis
FEM Thermal Model
Thermal
Analysis
Input
Material
Properties
Nodal Temperature
Distribution
Structural
Thermal
Loads
Component
Local Thermal
Environments
Engineering
Methods
(MINIVER)
Aeroheating
DownloadedbyKeiYunLauonNovember7,2016|http://arc.aiaa.org|DOI:10.2514/6.2012-5842
11. American Institute of Aeronautics and Astronautics
11
Figure 11, Lower Surface Peak Heating Results at 28 Seconds Mapped to Thermal Model
Figure 12, Aluminum Surface Panel Temperatures at 35 Seconds after Re-entry
The final thermal model removed the conservatism in the pseudo 3-D analysis discussed in Section II and
validated that the aluminum airframe provides sufficient thermal protection for the HIFiRE-4 mission. It defined the
internal subsystem compartment thermal environment to allow component level thermal analysis if necessary. More
importantly, it provided detailed internal temperature and temperature gradient distribution which can be mapped to
the finite element structural model (FEM) for structural analysis.
Time = 28Time = 28
Lower SurfaceUpper Surface
°F C°
450.0 232.2
425.3 218.5
400.7 204.8
376.0 191.1
351.3 177.4
326.7 163.7
302.0 150.0
277.3 136.3
252.7 122.6
228.0 108.9
203.3 95.2
178.7 81.5
154.0 67.8
129.3 54.1
104.7 40.4
80.0 26.7
DownloadedbyKeiYunLauonNovember7,2016|http://arc.aiaa.org|DOI:10.2514/6.2012-5842
12. American Institute of Aeronautics and Astronautics
12
Figure 13, Internal Cross Section Temperatures at 35 Seconds after Re-entry
IV. Structural Analysis of HIFiRE-4
The finite element structural model (FEM) was generated with quadratic tetrahedral elements for all components
including wing, fuselage/body, vertical fins, leading edges and control surfaces, shown in Figure 14. The model
contains 647,000 nodes and 330,000 elements. Most of the vehicle components were made of aluminum Al6061-
T6. Copper C11000 was used for the nose tip and leading edges of wing and body where high temperatures are
expected. To simplify the modeling, fuselage and leading edges were attached to other parts at their interfaces
through rigid connection elements simulating fasteners.
The temperatures for a flight trajectory were generated by transient thermal analysis with SINDA as described in
Section III. Temperatures at 35 sec after re-entry, considered as the maximum thermal loads condition, were mapped
to the structural model for strength analysis. The mapping between thermal to structural models was completed
utilizing a consistent finite element approach; a comparison is shown in Figure 15. Small differences can be seen
between the two models which are primarily due to the contrast of fringe colors used by different post-processing
tools, i.e., IDEAS vs. PATRAN. The geometry mismatch between thermal and structural models in local areas such
as nose tip and fins resulting in temperature extrapolations may have also contributed to the differences. These
differences are acceptable for the current analyses. The flight maneuver/aerodynamic loads were applied from an
Euler CFD solution at the maximum air loads condition (Mach 5.67, α=25°, Elevon -6.25°) as shown in Figure 16.
Linear strength analysis was performed for HIFiRE-4 with combined maximum aerodynamic pressure and
thermal loads. Temperature dependent material properties for both aluminum and copper were included in the
analysis. The maximum upward deflection with loads at the nose tip is 1.75 cm (0.689 in) and at the tail end is
about 0.439cm (0.173 in). The deformation fringe plots and vertical displacements along the centerline of the
bottom surface are illustrated in Figure 17. Von Mises stress distributions for both upper and lower surfaces of the
vehicle are shown in Figure 18 and are generally low, i.e., less than 155 MPa (22.5 ksi), except at a few areas close
to rigid connection elements that approximated the fastener installation and caused stress concentrations. A margin
check of the vehicle strength under ultimate loads has been conducted with the allowable strength for Al6061-T6
and C11000, in which the temperature dependency of allowable stress was considered. A few very small areas with
a negative margin were observed, which was judged to be a result of high temperature or high thermal gradients that
significantly reduce allowable stress in these areas. Stress concentration is very localized and a result of using RBE
connections to simulate fasteners. The high stresses are an artifact of modeling simplification and could be reduced
by including the contact modeling. The approximation is considered acceptable for the current study. The refined
analysis will be out of the scope of this low cost approach.
•Left Center Right Max Flipper
°F C°
450.0 232.2
425.3 218.5
400.7 204.8
376.0 191.1
351.3 177.4
326.7 163.7
302.0 150.0
277.3 136.3
252.7 122.6
228.0 108.9
203.3 95.2
178.7 81.5
154.0 67.8
129.3 54.1
104.7 40.4
80.0 26.7
DownloadedbyKeiYunLauonNovember7,2016|http://arc.aiaa.org|DOI:10.2514/6.2012-5842
13. American Institute of Aeronautics and Astronautics
13
Figure 14, HIFiRE-4 Finite Element Model Grids
Figure 15, Mapped Temperatures in Structure and Thermal Models at 35 Seconds.
Figure 16, Euler CFD Pressure Loads Mapped to FEM
DownloadedbyKeiYunLauonNovember7,2016|http://arc.aiaa.org|DOI:10.2514/6.2012-5842
14. American Institute of Aeronautics and Astronautics
14
Figure 17, Deflections of HIFiRE-4 under Maximum Loads
Figure 18, von Mises Stress due to Aerodynamic Pressure and Thermal Loads on HIFiRE-4
The original leading edge mounting was a tongue and groove design. The Australian DSTO partner was
concerned about the requirement in precision machining and potential problem of mismatched thermal growth. A
lap joint mounting was easier to fabricate and free of thermal mismatch.
A finite element analysis of contact modeling for the lap joint between leading edge (LE) and wing of vehicle
was performed with a model containing a small section of wing and leading edge, as shown in Figure 19. The LE
was attached to the wing with fasteners, and the wing section was constrained at all faces except the top and bottom
surfaces with uniform pressure applied at the lower LE surface, as illustrated in Figure 20. Two cases, one with and
one without contact surfaces between two components, were analyzed. A significant difference in deflections
between the two cases, Figure 21, indicated that the contact between LE and wing significantly reduces both
deflection and stress under pressure loads. Since the gap between the parts is very small, the potential contact
between them would limit the deformation of the LE and impose different boundary conditions that would result in a
different deformation for the LE. The no-contact model represents the conservative upper bound and showed that the
deflection of the leading edge would not exceed 0.01 inch. Contact of the leading edge with the wing will likely
limit the leading edge deflection to less than 0.0005 inch. Based on the above results and assumptions, the lap joint
leading edge design is not anticipated to have a critical structural integrity issue.
Deflection at Bottom Center Line
-0.1
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
0 20 40 60 80 100
X-location
Z-displacement
Leading edge deforms upward
by 0.689”, and tail deforms
upward by about .14”
Leading edge deforms upward
by 0.689”, and tail deforms
upward by about .14”
Lower Surface Upper SurfaceLower Surface Upper Surface
DownloadedbyKeiYunLauonNovember7,2016|http://arc.aiaa.org|DOI:10.2514/6.2012-5842
15. American Institute of Aeronautics and Astronautics
15
Figure 19, Wing Leading Edge Lap Joint FEM Submodel Location
Figure 20, Wing Leading Edge Lap Joint FEM Submodel Details
Section of interest
3.2psi Pressure Applied at
Bottom of Leading Edge
Bolts Connecting
Wing and Leading
Edge
•Wing Supported at
Three Faces
DownloadedbyKeiYunLauonNovember7,2016|http://arc.aiaa.org|DOI:10.2514/6.2012-5842
16. American Institute of Aeronautics and Astronautics
16
Figure 21, Deflections of Leading Edge with and without Contact Modeling
V. Conclusion
The analyses showed that a robust airframe can be designed using low cost materials and manufacturing process.
The design analyses were conducted with low labor budget. The feasibility study in Section 2 proved the
possibility of a heatsink airframe design. It allowed a detailed thermal/structural analysis that proved and
optimized the structural integrity in combined peak thermal and air loads to proceed.
1. Aluminum airframe design meets HIFiRE-4 mission requirements.
a. Lower surface skin temperatures approach aluminum’s one time use limit, but do not exceed.
b. No thermal coating would be required.
c. Airframe distortion has no adverse performance impact and is thermally induced. (Air loads distortion is
negligible.)
d. No excessive stress was predicted during flight.
2. Copper leading edge and aluminum tail fin design meets mission requirements.
3. Copper leading edge temperatures drop rapidly from stagnation to aluminum allowable at attachment points.
4. Current leading edge lap joint attachment design will meet mission requirements.
5. The MINVER engineering aeroheating code was shown to be very effective in full mission profile aerothermal
analysis.
6. The automatic processes of mapping aerothermal data to thermal model and thermal data to structural model
worked very well.
7. The CAD to detail 3-D thermal model is still labor intensive. The grid needs to capture both temperature
distribution and temperature gradients.
Acknowledgments
The successful completion of the HIFiRE-4 airframe design was due to the dedicated teamwork and hard work
that was performed by the Boeing thermal-structural team. The aeroheating and air loads analyses were supported by
Yuk K. Woo and Eric R. Unger. Preliminary thermal analysis were conducted by Ross D. Rochat and David S.
VanMiddendorp. Detail thermal analysis were conducted by John Q. Tran and Rowland W. Huang. The structural
analysis team included Richard R. Jacobs, Robert Quiroz, Aristidis Sidiropoulos, Jian P. Shao and Tsair-Jyh
(George) Tzong,
References
1
Douglas Dolvin, “Hypersonic International Flight Research and Experimentation (HIFiRE) Fundamental Science
and Technology Development Strategy”, AIAA-2008-2581, 15th AIAA International Space Planes and
Hypersonic Systems and Technologies Conference, Dayton, Ohio, Apr. 28-May 1, 2008.
2
Thomas R. Smith, Kevin G. Bowcutt, etc., “HIFiRE 4: A Low-Cost Aerodynamics, Stability, and Control
Hypersonic Flight Experiment”, AIAA-2011-2275, 17th AIAA International Space Planes and Hypersonic
Systems and Technologies Conference, San Francisco, California, Apr11-14, 2011
3
Joseph M. Hank, James S. Murphy, and Richard C. Mutzman, “The X-51A Scramjet Engine Flight
Demonstration Program”, AIAA Paper 2008-2540, 15th AIAA International Space Planes and Hypersonic
Systems and Technologies Conference, 28 April - 1 May 2008, Dayton, Ohio.
DownloadedbyKeiYunLauonNovember7,2016|http://arc.aiaa.org|DOI:10.2514/6.2012-5842
17. American Institute of Aeronautics and Astronautics
17
4
Laurie A. Marshall, Griffin P. Corpening and Robert Sherrill, “A Chief Engineer's View of the NASA X-43A
Scramjet Flight Test”, AIAA Paper 2005-3332, 13th AIAA International Space Planes and Hypersonic Systems
and Technologies Conference, Capua, Italy, May 16-20, 2005.
5
Kimmel, R. L., Adamczak, D., Gaitonde, D., Rougeux, A., Hayes, J. R., “HIFiRE-1 Boundary Layer Transition
Experiment Design,” AIAA paper 2007-0534, January 2007.
6
David Manor, Kei Y. Lau and Donald B. Johnson, “Aerothermodynamic Environments and Thermal Protection
for a Wave-Rider Second Stage”, Journal of Spacecraft abd Rockets, Vol. 41, No. 4, July-August 2004.
7
C. D. Engle and S. C. Prahara, MINIVER Upgrade for the AVID System, NASA CR-172212, 172213, 172214,
August 1983.
8
K. Y. Lau, L. F. Scuderi, D. H. Petley and P. W. Yarrington, Dual Fuel Vehicle Thermal Management, AIAA
7th International Space Planes Conference, Norfolk VA, November 1996, AIAA paper 96-4595.
9
K. E. Wurster, E. V. Zoby and R. A. Thompson, Flowfield and Vehicle Parameter Influence on Results of
Engineering Aerothermal Methods, Journal of Spacecraft and Rockets, Vol 28, No. 1, Jan-Feb 1991, pp. 16-22.
10
K. Y. Lau and Edward A. Eiswirth, “Hypersonic Boundary Layer Transition – Application to High Speed
Vehicle Design”, Invited AIAA Paper 2007-0310, 45th
AIAA Aerospace Sciences Meeting, Reno NV, January
2007.
DownloadedbyKeiYunLauonNovember7,2016|http://arc.aiaa.org|DOI:10.2514/6.2012-5842
18. This article has been cited by:
1. Bianca R. Capra, L. M. Brown, R. R. Boyce, S. C. Tirtey. 2015. Aerothermal–Structural Analysis of a Rocket-Launched Mach
8 Scramjet Experiment: Ascent. Journal of Spacecraft and Rockets 52:3, 684-696. [Abstract] [Full Text] [PDF] [PDF Plus]
DownloadedbyKeiYunLauonNovember7,2016|http://arc.aiaa.org|DOI:10.2514/6.2012-5842
