17th AIAA International Space Planes and Hypersonic Systems and Technologies Conference
11 - 14 April 2011, San Francisco, California
AIAA 2011-2275
HIFiRE 4: A Low-Cost Aerodynamics, Stability, and Control
Hypersonic Flight Experiment
Thomas R. Smith1, Kevin G. Bowcutt2, John R. Selmon3, Luis Miranda4, Brandin Northrop5, Ron Mairs6, and Eric
R. Unger7
The Boeing Company, Huntington Beach, CA 92647
Kei Y. Lau8
The Boeing Company, Berkeley, MO 63134
Todd Silvester9, Hans Alesi10, Allan Paull11, and Ross Paull 12
Defence Science and Technology Organisation - Brisbane, Kenmore, QLD, Australia 4069
and
Douglas J. Dolvin13
AFRL/RBAH, Wright Patterson AFB, OH 45433
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). Boeing is an industrial
partner in the HIFiRE program, having joint responsibility with DSTO for three flight
experiment payloads (numbers four, seven, and eight). The overall objective of the HIFiRE
program is to gather fundamental scientific and engineering data on the physics and
technologies critical to future operational hypersonic flight by conducting several flight
experiments using a relatively low cost sounding rocket based flight test technique advanced
by the co-authors from DSTO. The specific objective of HIFiRE flight experiment number
four (HIFiRE 4), the subject of this paper, is to gather flight data on the aerodynamics,
stability, and control of an advanced waverider configuration. This flight is also a precursor
to HIFiRE flight experiment number eight (HIFiRE 8), which will integrate an advanced
three-dimensional scramjet engine with the HIFiRE 4 airframe to flight test the combined
aero-propulsive configuration. Discussed is the design and performance of HIFiRE 4, with
emphasis given to vehicle guidance and control.
1
Associate Technical Fellow, Platform Performance Technology, 5301 Bolsa Ave./H45N-E408, AIAA Senior
Member.
2
Senior Technical Fellow & Chief Scientist of Hypersonics, Platform Performance Technology, 5301 Bolsa
Ave./H45N-E406, AIAA Fellow.
3
Senior Engineer/Scientist, Vehicle Management Systems Technology, 5301 Bolsa Ave./H45N-E408.
4
Engineer/Scientist, Vehicle Management Systems Technology, 5301 Bolsa Ave./H45N-E408, AIAA Senior
Member.
5
Engineer/Scientist, Platform Performance Technology, 5301 Bolsa Ave./H45N-E408.
6
Senior Engineer/Scientist, Platform Performance Technology, 5301 Bolsa Ave./H45N-E408.
7
Senior Engineer/Scientist, Platform Performance Technology, 5301 Bolsa Ave./H45N-E408.
8
Technical Fellow, Aero and Flight Controls, 5621 James S McDonnell Blvd./S064-7640, AIAA Senior Member.
9
Senior Aerostructures Engineer, Applied Hypersonics Branch, Air Vehicles Division, PO Box 883.
10
Principal Aerostructures Engineer, Applied Hypersonics Branch, Air Vehicles Division, PO Box 883.
11
Research Leader, Applied Hypersonics Branch, Air Vehicles Division, PO Box 883.
12
Principal Systems Engineer, Applied Hypersonics Branch, Air Vehicles Division, PO Box 883.
13
Program Manager, Aerospace Flight Research and Experimentation, 2130 Eighth St., Ste 1, Area B/BLDG 45/RM
256, AIAA Senior Member.
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Copyright © 2011 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.
Nomenclature
α
β
δs
M
CL
CD
L/D
Cm
Cp
=
=
=
=
=
=
=
=
=
angle of attack
angle of sideslip
control surface deflection angle
Mach number
vehicle lift coefficient
vehicle drag coefficient
lift to drag ratio
moment coefficient
pressure coefficient
I. Introduction
T
he 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). Boeing is an industrial partner in the HIFiRE program, having joint responsibility with DSTO
for three flight experiment payloads (numbers four, seven, and eight). The overall objective of the HIFiRE program
is to gather fundamental scientific and engineering data on the physics and technologies critical to future operational
hypersonic flight by conducting several flight experiments using a relatively low cost sounding rocket based flight
test technique advanced by the co-authors from DSTO. The primary objective of HIFiRE 4 is to partially pull out
from a near vertical atmospheric re-entry by a nominal altitude of 27km while remaining under-control at the endpoint. The extent of the pull-out is not critical, however, it is expected the partial pullout will result in a 20o
alteration from a ballistic trajectory. The secondary objective is to then bring the vehicle to the ground via 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. This flight is also a precursor to HIFiRE flight experiment number eight (HIFiRE 8), which will integrate an
advanced three-dimensional scramjet engine with the HIFiRE 4 airframe to flight test the combined aero-propulsive
configuration.
Requirements for designing the HIFiRE 4 vehicle include aerodynamic performance, stability, and control
adequate for a high angle-of-attack pull-up maneuver from a flight path angle of about negative 70-degrees to zero
(i.e., horizontal); high lift-to-drag ratio during low angle-of-attack hypersonic flight; adequate internal volume and
volume distribution to house all required subsystems and instrumentation; structural materials suitable for the aerothermal environment encountered in flight; a configuration that can accommodate scramjet integration; and a mass
small enough that it can be boosted to Mach 6-8 with a VSB-30 sounding rocket stack. To accommodate these
requirements a configuration was developed around an osculating cone waverider1, 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 1.
[1] Sobieczky, H., Dougherty, F. C., and Jones, K. D., "Hypersonic Waverider Design from Given Shock Waves," in
Proceedings of the First International Hypersonic Waverider Symposium, University of Maryland, College Park,
MD, Oct. 17-19, 1990.
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2154 mm
512 mm
289 mm
Figure 1. Three-view of the HIFiRE 4 flight experiment vehicle illustrates the truncated osculating cone
waverider shape used for its design.
The HIFiRE 4 vehicle will be constructed of conventional materials: aluminum for the body and copper for the
leading edges. These relatively low-temperature materials are viable because of the trajectory flown by the vehicle,
which 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, keeping the airframe integrated heat load low enough to permit the use of aluminum
and copper as structural materials. Aerodynamic heating and thermal analysis results will be presented to validate
this choice of airframe materials.
The HIFiRE 4 two-stage rocket spins up after launch via canted fins in order to stabilize the attitude and
trajectory of the unguided launch stack and therefore reduce trajectory dispersions. After exiting the atmosphere the
rocket second stage and payload will be de-spun, the payload shroud will be deployed, two HIFiRE 4 vehicles will
be released from the second stage structure and cold-gas nitrogen thrusters will be used to align the attitude of each
vehicle for atmospheric entry. The same exo-atmospheric control strategy will be utilized on both flyers. Following
atmospheric entry and dynamic pressure building up to a level adequate for aerodynamic control, vehicle control
will transition from gas thrusters (exo-atmospheric control) to elevons (endo-atmospheric control) for each vehicle.
The two vehicles will be identical in configuration but will be controlled at this stage using different control
algorithms. One algorithm will be developed by Boeing; the other will be developed by DSTO. After the partial
pull-up maneuver is complete, one vehicle will continue to be controlled to horizontal ground impact at a low
subsonic speed. The other will undertake a controlled maneuver to provide additional information on stability. Flight
trajectories and flight control simulations will be presented to demonstrate vehicle stability and control, and the
ability to successfully guide and control the HIFiRE 4 vehicle along its desired trajectory with sufficient margins to
accommodate uncertainties.
The HIFiRE 4 vehicle employs a control system comprised of a roll-stabilized inertial measurement unit (IMU),
GPS, two control elevons, control actuators, a nitrogen tank and cold gas thrusters, and a flight computer. Other
subsystems include batteries for electrical power, servo control boards, power switching boards, sensors to measure
pressure and temperature on the outside and inside of the vehicle, and a telemetry system to transmit all desired
pressure, temperature, IMU, and flight computer data to ground receiving stations. All vehicle subsystems are
illustrated in a preliminary internal arrangement in Figure 2.
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N2 Tank
for Cold Gas
Thrusters
Cold Gas Thrusters
Linear
Actuators
IMU
Pressure & Temp Sensors
Batteries
Magnetometer
Tungsten
Ballast
Copper
Leading Edges
Transponder and
Computer
Transmitter and
Computer
Aluminum Body
Figure 2. The internal arrangement for HIFiRE 4 illustrates all subsystems and their placement within the
vehicle.
The HIFiRE 4 experimental payload is scheduled to fly at the Woomera Test Range in Australia in October of
2012. Preliminary design of the HIFiRE 4 vehicle is nearly complete. The status of vehicle design and analysis will
be presented, as will a conceptual view of how the HIFiRE 7 scramjet with an advanced three-dimensional inlet
might be integrated on the HIFiRE 4 airframe to become HIFiRE 8.
II. Requirements
The HIFiRE 4 program has four sets of requirements: scientific, experimental, system, and vehicle requirements.
The following sections define these requirements and provide the mission description.
A. Scientific Requirements
• Gather flight data on the aerodynamic performance and control of an advanced hypersonic
configuration.
• Validate hypersonic vehicle design and analysis tools.
B. Experimental Requirements
Experimental Purpose
• Develop and validate booster separation and attitude control strategies for a hypersonic boost-glide
vehicle.
• The control strategy must provide for both exo-atmospheric and endo-atmospheric vehicle control.
Primary Experiment Requirements
• Demonstrate exo-atmospheric vehicle stack de-spin.
• Demonstrate exo-atmospheric nose cone and then vehicle separation.
• Demonstrate exo-atmospheric vehicle attitude control to align vehicle axis with velocity vector.
• Demonstrate vehicle attitude control to 25degree AOA at atmospheric entry point.
• Alter ballistic entry trajectory via partial pull-up that terminates at a nominal altitude of 27km while
maintaining vehicle control.
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Secondary Experiment Requirements
The secondary experiments begin after partial pull-up. The secondary experiment for the DSTO controlled
vehicle remains to be defined. The vehicle being controlled by Boeing, however, will endeavor to:
• Continue pull-up to horizontal flight and then “land” (i.e., impact the ground at a slow speed and low
rate of decent)
Tertiary Experiment Requirements
• Allowances have been made for the inclusion of small (non-instrumented) material test panels on the
underside of the vehicle.
C. System Requirements
• Exo-atmospheric de-spin system.
• Exo-atmospheric nosecone and flyer separation system.
• Exo-atmospheric attitude sensing and control system.
• Endo-atmospheric GNC system, including control effectors, actuators, attitude sensing relative to the
freestream flow, and a flight control strategy and algorithm.
D. Vehicle Requirements
• Dynamically stable with automated flight control.
• Stable in yaw when flying in the atmosphere.
• Mach 8 maximum speed.
• Flyer to fit within a 640-mm diameter payload shroud.
• Length approximately 2 m.
• Flyer mass approximately 100 kg.
• Exo- and endo-atmopheric control effectors.
• Sufficient fuselage volume and form factor to package all required hardware.
• No exotic materials or complex-to-fabricate parts.
E. Mission Description
The HIFiRE 4 flyers will be launched together on a VSB-30 sounding rocket. The mission profile is shown in
Figure 3 and the mission events are listed in Table 1. The rocket stack is spin stabilized. At an altitude of 100 km a
yo-yo de-spin system will deploy and reduce stack spin rate. At 120 km the two flyers will be released. After
approximately 70 s from launch an exo-atmospheric control maneuver will commence. At this stage both flyers are
controlled independently, but utilize the same control strategy. Cold-gas thrusters will be used to reorient the
vehicles in roll, pitch and yaw prior to re-entry. The hand-over point from exo- to endo-atmosphere control will take
place at an atmospheric dynamic pressure adequate for aerodynamic control. At this point both vehicles will possess
a 25o angle of attack and the control system will trim the vehicle at this attitude during the pull-up. The primary
experiments will end at an altitude of approximately 28km. The DSTO flyer then falls to an impact in the Woomera
Test Range. At 28 km the Boeing secondary experiment begins, whereby the vehicle will execute a complete pull-up
to horizontal flight and before continuing under controlled to a low-subsonic speed, low rate of decent ground
impact.
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Figure 3. HIFiRE 4 mission description.
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Table 1. HIFiRE 4 Mission Events.
No.
Mission Event
Time
Alt.
Mach
[sec]
[km]
No.
a
0
S31 ignition is initiated by FCU.
T +0
b
0
0
1
S31 motor burns out.
T+12
4
1.9
2
S31 separates from S30.
T+13.5
4.5
1.9
3
S30 motor ignition is initiated by timer.
T+16.6
5.9
1.9
4
S30 motor burns out.
T+45.5
41
7.3
5
Yo-Yo de-spin system is initiated by timer.
T+60
100
7.9
6
Nosecone separation is initiated by timer.
T+65
112
7.1
7
Payload separation is initiated by timer.
T+68
120
5.7
8
Exo-atmosphere attitude control maneuvers for DSTO and Boeing
T+70
125
-
free-flyers commence.
9
Free-flyers coast through trajectory apogee.
T+292
292
-
10
Exo-atmosphere attitude control maneuvers end.
T+520
112
7.1
11
Configurations re-enter the upper atmosphere at an altitude of 100
T+525
100
7.9
km.
12
In-atmosphere attitude control maneuvers begin.
T+530
88
7.9
13
DSTO pull-up experiment ends. Boeing experiment begins.
T+558
29.5
6.4
14
DSTO payload impacts.
T+590
0
0.3
15
Boeing pull-up complete.
T+565
24.6
3.9
16
Boeing payload lands.
T+1577
0
0.19
III. Vehicle Configuration
The general arrangement of the HIFiRE 4 flight experiment is shown in Figure 4. The concept consists of a
truncated waverider wing with wing tip vertical stabilizers and a fuselage on top of the wing. The waverider wing
and stabilizers have 1.0 mm leading edge radii. The vertical stabilizers hang down from the wing and the leading
edge of the verticals rides on the waverider shockwave at Mach 8 to retain compression produced by the shockwave.
The thermal approach is a thick heat sink primary aluminum structure with copper leading edges to pull the heat
away form the sharp leading edge radius quickly and transfer it to the aluminum. The wing is an aluminum slab with
a long central pocket down the center to hold internal components. This simple approach to fabrication the wing also
reduces cost. There are long pockets cut top and bottom to house pressure transducer instrumentation. Removable
aluminum panels cover the instrumentation pockets.
The leading edge of the wing and vertical stabilizers are made of copper which is segmented to allow differential
thermal expansion. Copper was chosen for its high thermal conductivity. The fuselage has a “bread-loaf” crosssection. The forward half is ogival and tapers into the wing at the leading edge. The aft half is constant in crosssection. The skin of the fuselage is divided into three removable covers for access to the internal components. The
HIFiRE 4 flight experiment mass is 74.19 kg (163.5 lbs) dry and 74.87 kg (165.0 lbs) wet (i.e., with nitrogen RCS
propellant).
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Sensor/Thruster
Module
Elevons
Instrumentation Covers
HF-4 Data Summary
Structure
74.36 kg 163.9 lbs
Systems
18.74 kg
41.3 lbs
Dry Weight
93.10 kg 205.2 lbs
GN2
0.68 kg
1.5 lbs
Launch Weight
93.78 kg 206.7 lbs
Reference length
1.989 m
78.3 in
Xcg Station
1.166 m
45.9 in
Wing Reference Area
0.78 m2 8.36 ft2
Removable
Fuselage
Covers
Aluminum
Fins
Copper
Leading
Aluminum Wing
Pitot Sensor
Copper Leading
Edge
Figure 4. The isometric view of the HIFiRE 4 vehicle illustrates the external features.
IV. Structures
The structural concept is a one-piece wing machined from 6061-T6 aluminum plate with thinner removable
covers fastened to the wing. The leading edges are segmented C11000 copper components. The simplicity of the
concept is similar to wind tunnel model practices, reducing manufacturing costs. The extra weight incurred by not
having a built up wing is within booster performance capability. The wing aluminum bulk also serves the additional
function of being a heat sink to absorb thermal loads.
Figure 5 shows an exploded view of the HIFiRE 4 structure concept. All pieces that are not leading edge are
made of the same 6061-T6 material. The concept shown is preliminary. The final detail design will be performed by
DSTO in 2011. The vertical fins will likely be one-piece while the elevons will likely be built-up structure. The
covers will have a nominal 5 mm thickness.
The leading edges are made of C11000 Electrolytic Tough Pitch (ETP) copper. This material was chosen for its
high thermal conductivity. The sharp (1.0 mm) leading edges will experience a large heat flux at Mach 8. The
wedge-like copper leading edges conduct heat into the bulky heat-sink aluminum wing. The leading edges are
segmented to reduce stress due to differential thermal expansion. Boeing has completed detail design of the leading
edges and has procured two shipsets as of this writing.
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Sensor/Thruster
Module
Dorsal
Fuselage Covers
Instrument Covers
Elevons
Leading Edges
Verticals
Instrument Covers
Wing One Piece 6061-T6 plate
Figure 5. Exploded view shows HIFiRE 4 structure.
V. Systems
A. Flight Controls
1. Aerodynamic Flight Controls
The elevons will be controlled by electro-mechanical actuators. Utilizing the HIFiRE 4 guidance and control
simulation results, and generic motor characteristics, an actuator sizing analysis was performed. The analysis was
performed on the portside control surface so that the analysis would include additional load due to sideslip.
Moreover, and additional 10 deg of deflection was added to the trajectory results to account for differential
deflection that may be required for maneuvering. Finally, a 25% margin was added to the results. Figure 6 illustrates
the resulting load history. The maximum hinge moment is about 200 N-m. In order to estimate the power
requirements, a 28 volt power source was assumed. Figure 7 illustrates the analysis results for the portside elevon,
which shows a power consumption of 200 W-min in 1200 seconds.
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-50
-443
-100
-885
-150
symmetric trajectory
trajectory with 5° slideslip
throughout descent
trajectory with 10° of differential -1328
deflection
trajectory with 5° slideslip + 10°
of differential deflection
-200
-1770
-250
0
0.5
1
Time (Min.)
1.5
Figure 6. Actuator Sizing Loads.
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-2213
2
in-lbs
0
N-M
0
Port Side Current (Amps)
Port Side Power (KW)
Port Side Energy (KW-minutes)
8
6
4
2
0
0
200
400
600
800
1000
1200
1400
0
200
400
600
800
1000
1200
1400
0
200
400
600
800
1000
1200
1400
0.2
0.15
0.1
0.05
0
0.25
0.2
0.15
0.1
0.05
0
sim time (seconds)
Figure 7: Actuator Power Requirements.
2. Exo-atmospheric Flight Controls
The exo-atmospheric flight control system consists of a set of gaseous nitrogen (GN2) cold gas thrusters,
providing flight control at altitudes above ~60 km, ensuring that the flyers are oriented in the correct attitude when
they reenter the atmosphere. The GN2 will be supplied by a Luxfor brand aqualung gas cylinder that consumes
much of the fuselage volume (see Figure 2). Figure 8 shows the combined attitude sensing module and cold gas
thruster pod on the aft end of the flyer. The cylindrical pod has four thruster nozzles.
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3 x thruster block each side
IMU housing
Figure 8. The thruster block/IMU housing showing the original 6 thrusters proposed, which has now been
reduced to 4.
B. Avionics
The internal arrangement concept of the HIFiRE 4 flight experiment is shown Figure 2. The majority of
components are stored in the central fuselage cavity. The foremost component in the fuselage cavity is tungsten
ballast. Batteries are lithium-ion D-cell contained in a custom package to fit into the nose cavity. Each flyer will
house 5 DSP flight computers developed by DSTO based on a PCI-104 format. Servo control, power switching and
communication boards share the same footprint enabling them to be mounted in a stack arrangement as shown in
Figure 9.
The internal arrangement shown is notional. The final internal arrangement design will be performed by DSTO.
The complete set of hardware is listed below.
• 160 channels per flyer, distributed approximately as:
o 50 control and guidance sensors
o 64 pressure sensors
o 36 RTDs (thin film platinum temperature sensors)
• 5 DSP flight computers
• 1 power servo board (for elevon actuators)
• 1 power servo and inverter board to give 96V power supply
• 4 battery packs wired in parallel
• 1 power switching board
• DMARS-R IMU
• 1 pitot tube
• 1 magnetometer
• 1 DSTO accelerometer pack
• Combined GPS and TM wrap around antenna
• GPS receiver and GPS LNA
• 1 TM transmitter
• 1 TM conditioning board
• 2 RF switches (to switch GPS and TM from payload service module telemetry to flyer telemetry, after
shroud deployment)
• 4 gas thrusters for exo-atmospheric attitude control
• N2 reservoir for thrusters
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•
•
•
•
•
N2 temperature sensor
N2 pressure sensor
2 servo motors for elevon actuation
1 wide angle camera
Analogue conditioning boards
Controller Board
Mark I Flight Computer used
on HIFiRE 0 and 1
Power Switching Board
IBM PCI-104 DSP
Communications Board
Figure 9. Typical HIFiRE computer stack.
C. Sensors
HIFiRE 4 has two types of sensors: flight control sensors and experiment instrumentation. The flight sensors
include GPS, IMU, accelerometer packs, a flush pitot sensor on the vehicle nose, and a magnetometer. The IMU is
stored in the Thruster Block/IMU Housing shown in Figure 8. The GPS and magnatometer are located with the
flight computers. The pitot sensor is attached to the back of the foremost copper leading edge. Figure 10 illustrates
how the leading edge has a 1mm diameter hole machined into it to serve as a port for the pitot sensor.
The instrumentation includes 64 Honeywell SDX pressure transducers and 36 thin film RTD platinum
temperature sensors. The pressure transducers are stored in instrumentation cavities outboard of the fuselage in the
upper and lower surfaces of the wing. Figure 11 shows a typical installation where pressure transducers are mounted
to the instrumentation cavity covers. Pressure is sensed down the length of the vehicle, top and bottom on both
wings. Temperature sensors are located on the covers of the same cavities.
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2mm OD x 1.5mm ID and 1.4mm
OD x 1mm ID stainless steel tubing
1mm hole (EDM)
6.35mm OD x 2.1mm ID copper
spacer (57mm long. nom)
Figure 10. Flush pitot sensor installation.
Surface
mounted
pressure sensors
Hypodermic tubing
Resistive Thermal
Devices (RTDs)
mounted this surface
(not shown)
Figure 11. Typical pressure transducer installation.
VI. Booster and Booster Integration
The HIFiRE 4 flyers will be launched on a VSB-30 sounding rocket that was developed by the Institute of
Aeronautics and Space (IAE) of the Brazilian Air Force (FAB) for the German Aerospace Center (DLR). The VSB30 is shown in Figure 12.
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Figure 12. The VSB-30 sounding rocket.
The HIFiRE 4 flyers will be housed in a 750 mm diameter fairing illustrated in Figure 13. The cylindrical IMU
housings on the back of the two flyers slip into ejection cylinders. Tanks in the Payload Support Module (PSM)
supply pressurized gas to expel the flyers.
Ballast
Weight
(100kg)
Nosecone Tip
Carbon-Phenolic
Boeing Glider
Nosecone
Snubber
DSTO Glider
Nosecone
Manacle Band
PSM
Push-Off Cylinders
PSM
And Pistons
Extension
Figure 13. Two view of the fairing and deployment concept.
VII. Mass Properties
A bottoms-up approach based on the level of configuration detail available was used to calculate the preliminary
mass properties for the HIFiRE 4 flight vehicle. CAD models were used to calculate the mass, center of gravity, and
the moments of inertia of major structural components. Subsystem weights were built up from the component level
with appropriate allowances given for installation, wiring, and fastener weights (10%, 20% and 5%, respectively).
Table 2 shows the preliminary HIFiRE 4 mass properties. The longitudinal nose station is located at 0.0276 m in the
coordinate system used to define the location of all components.
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Table 2 Preliminary HIFiRE 4 mass properties.
Mass Moment of Inertia about
Center Of Gravity
HiFire CG
X-cg
Y-cg
Z-cg
Ixx
Iyy
Izz
Group
Mass
Mass
kg
lbs
m
m
m
kg-m2
kg-m2
kg-m2
Structure
74.36
163.9
1.10
0.00
-0.07
Actuators
6.70
14.8
1.90
0.00
-0.11
Avionics
6.51
14.3
0.82
0.00
-0.03
Propulsion
3.69
8.1
1.66
0.00
-0.07
Instrumentation
1.84
4.1
1.27
0.00
-0.08
Total Dry
GN2
Total Wet
93.10
0.68
93.78
205.2
1.5
206.7
1.164
0.000
-0.071
1.51
0.00
-0.07
1.166
0.000
-0.071
Mass Product of Inertia about
HiFire CG
Pxy
Pxz
Pyz
kg-m2
kg-m2
kg-m2
1.82
30.00
31.25
0.00
-2.41
0.00
1.82
30.08
31.33
0.00
0.00
0.00
VIII. Aerodynamic Analysis
Figure 14. CFD analysis configuration.
A. CFD Modeling
The HIFiRE 4 configuration was analyzed using the Cart3D CFD code to predict inviscid forces and
moments. The Blasius equation with a compressibility correction was used to predict laminar friction drag and the
van Driest II method was used to predict turbulent friction drag. The Euler plus friction drag approach was validated
by comparison with RANS CFD results and will be discussed below. Figure 14 shows the configuration that was
analyzed. The Cart3D code is an Euler (inviscid), cell-centered, finite-volume, upwind-differencing, computational
fluid dynamics solver. It has been developed to capture wave propagation with very low dissipation. The tool is used
extensively by NASA and is applicable from transonic to high supersonic Mach numbers. The power behind Cart3D
lies in its utilization of cartesian volume meshes that are built automatically around a triangular surface definition of
the geometry, thus greatly minimizing a user’s time generating a grid. In a further step towards streamlining,
individual components of a given configuration may be defined separately and later intersected using tools within
the Cart3D suite. One application where this ability is especially powerful is for movable components, such as fins.
The user merely needs to rotate a given component, and Cart3D handles the rest, with no need to recreate any
surface meshes. This capability was utilized extensively within this program, where desired flap deflections were
accommodated by using external scripts to rotate the individual flaps and Cart3D handled the rest of the workload in
an automated fashion.
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B. CFD Results
Early IRAD-funded CFD investigations of the HIFiRE 4 vehicle were focused on longitudinal data along
with a limited set of lateral-direction analysis points for the high angle of attack entry portion of the flight trajectory.
The data points covered a range of Mach number, angle of attack, angle of sideslip, symmetrical flap deflections,
and asymmetrical flap deflections (the latter at high Mach numbers only). Within the various narrow flight
envelopes defined, a fully factored database was generated where needed. In follow-on work we significantly
expanded the database to include a much larger flight envelope with more extensive lateral-directional data. While
the initial decision to define symmetrical and asymmetrical flaps settings seemed a good approach to distinguish
control for pitch and roll separately, it was soon deemed unwieldy for the much larger database. It also was realized
that a fully factored database in all flight parameters, including flap deflections, would be extremely large (about
20,000 CFD analysis point) and therefore not possible within the scope of the program. A decision was therefore
made to treat the two flaps is if they behave in a decoupled fashion by assuming there are no interference effects
between the port and starboard flaps. Using this approach only a single flap need be moved to populate the entire
database (with the opposite flap held fixed at zero deflection). In high Mach flow this is a very good assumption, but
testing was needed to quantify the errors at low Mach numbers. The parameters defining the full database are
detailed in Table 3. Note that data for the effects of opposite flap deflections are determined using symmetry
arguments.
Table 3: Aerodynamic Database Analysis Points.
Mach
α
β
0.3, 0.5, 0.8, 0.95, 1.05
1.2, 1.5, 2.0, 3.0, 4.0
5.0
6.0, 8.0
-5, 0, 5, 10, 15, 20, 25, 30
-5, 0, 5, 10, 15, 20, 25, 30
5, 10, 15, 20, 25, 30
20, 25, 30
-5, -2.5, 0, 2.5, 5
-5, -2.5, 0, 2.5, 5
-5, -2.5, 0, 2.5, 5
-5, -2.5, 0, 2.5, 5
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δs
-20, -15, -10, -5, 0, 5
-20, -15, -10, -5, 0, 5
-20, -15, -10, -5, 0, 5
-20, -15, -10, -5, 0, 5
Plotted in Figure 15 are inviscid longitudinal aerodynamic force and moment coefficients at Mach 7, and
Figure 16 shows the same data for Mach 2. These data indicate that the vehicle is generally well behaved in terms of
stability, but slightly unstable in pitch over much of the conditions shown. A similar series of plots for lateraldirectional data at Mach 2 are shown in Figure 17. Note that while yawing moment plot shown indicates that the
vehicle is stable for these particular flight conditions, this was not true for a large portion of the flight envelope,
particularly at lower-speed conditions. To help resolve this problem a dorsal fin was added (as shown in Figure 14).
A sample of one solution with the dorsal fin can be seen in Figure 18 for Mach 6, zero angle of attack, zero side-slip
and elevons deflected -5°. Generally, the added dorsal fin has a marginal affect on yaw stability and is effective only
at negative angles-of-attack, as expected, necessitating a roll maneuver to invert the vehicle for low-speed flight.
As was mentioned, viscous effects were modeled using the Blasius equation with a compressibility correction
to predict laminar friction drag and the van Driest II method to predict turbulent friction drag. To validate the Euler
plus friction drag approach we ran the Overflow RANS CFD code at Mach 7 and 115 kft altitude to generate viscous
solutions to compare with Euler CFD plus friction drag calculated using the flat plate methods specified. Figure 19
shows an example of Euler plus friction drag predictions at Mach 7 and 115 kft. Figure 20 shows a comparison
between Cart3D and Overflow solutions run at Mach 7, 115 kft altitude, 25 deg angle of attack, zero side-slip, and
elevons deflected -5 deg. The pressure plots compare well. Figure 21 then presents comparisons of Euler plus
friction vs. RANS for Mach 7. The lift curve, drag polar, and moment curve shown indicate an excellent match
between the two methods.
Mach 7 Lift Curves
0.5000
0.7000
0.4500
0.6000
0.4000
0.5000
CL
delta=-15
CD
delta=-20
0.4000
0.3500
delta=-15
0.3000
delta=-20
0.2500
delta=-10
0.3000
delta=-10
0.2000
delta=-5
0.2000
delta=-5
delta=0
0.1500
delta=0
delta=5
0.1000
delta=5
0.1000
0.0000
-0.1000
-10.00
0.0500
0.00
10.00
20.00
0.0000
-0.10
30.00
0.00
0.10
0.20
Alpha
0.30
0.40
0.50
0.60
Mach 7 L/D Curves
10.0
0.0200
8.0
0.0150
delta=-20
delta=-15
delta=-20
6.0
0.0100
delta=-10
0.0050
delta=-5
4.0
delta=-10
2.0
delta=-5
Cm
L/D
delta=-15
delta=0
0.0000
delta=5
delta=0
0.0
delta=5
-0.0050
-2.0
-0.0100
-4.0
-0.0150
-6.0
-10.00
0.00
10.00
0.70
CL
20.00
30.00
-0.0200
-10.00
0.00
10.00
Alpha
Alpha
Figure 15. Longitudinal data at Mach 7.
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20.00
30.00
Mach 7 Lift Curves
0.7000
0.5000
0.6000
0.4500
0.4000
0.5000
delta=-20
delta=-15
CD
CL
0.4000
0.3500
delta=-15
0.3000
delta=-20
0.2500
delta=-10
0.3000
delta=-10
0.2000
delta=-5
0.2000
delta=-5
delta=0
0.1500
delta=0
delta=5
0.1000
delta=5
0.1000
0.0500
0.0000
-0.1000
-10.00
0.00
10.00
20.00
0.0000
-0.10
30.00
10.0
0.10
0.20
0.30
0.40
0.50
0.60
0.70
0.0400
8.0
delta=-20
0.0300
delta=-20
6.0
delta=-10
2.0
delta=-5
delta=-5
delta=0
delta=5
0.0000
delta=0
0.0
delta=-10
0.0100
Cm
4.0
delta=-15
0.0200
delta=-15
L/D
0.00
CL
MachAlpha
7 L/D Curves
delta=5
-0.0100
-2.0
-0.0200
-4.0
-6.0
-10.00
0.00
10.00
20.00
-0.0300
-10.00
30.00
0.00
10.00
Alpha
20.00
30.00
Alpha
Figure 16. Longitudinal data at Mach 2.
0.0250
0.0030
0.0200
delta=-20
0.0150
delta=-15
0.0100
delta=-10
CY
delta=0
0.0000
Cl
delta=-5
0.0050
delta=-20
0.0025
delta=5
-0.0050
delta=-15
0.0020
delta=-10
0.0015
delta=-5
delta=0
0.0010
delta=5
0.0005
-0.0100
0.0000
-0.0150
-0.0005
-0.0200
-0.0010
-0.0250
-6
-5
-4
-3
-2
-1
0
1
2
3
4
5
-6
6
-5
-4
-3
-2
-1
Beta
1
2
3
4
5
6
0.0050
0.0060
0.0040
delta=-20
0.0040
delta=-20
0.0020
delta=-15
0.0030
delta=-15
0.0000
delta=-10
0.0020
delta=-10
delta=-5
delta=0
-0.0040
delta=5
delta=0
0.0000
-0.0060
-0.0010
-0.0080
-0.0020
-0.0100
-0.0030
-0.0120
-0.0040
-0.0140
delta=-5
0.0010
Cm
-0.0020
Cm
0
Beta
delta=5
-0.0050
-6
-5
-4
-3
-2
-1
0
1
2
3
4
5
6
-6
-5
-4
-3
-2
Beta
-1
0
Beta
Figure 17. Lateral-directional data at Mach 2.
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1
2
3
4
5
6
Lower
Surface
Upper
Surface
Figure 18. Pressure Coefficients on Vehicle at M=6, α=0, δ=-5° & β=0.
6
0.45
5
0.40
0.35
4
0.30
CD
L/D
3
2
1
0.25
0.20
0.15
0
0.10
-1
0.05
0.00
-2
-10
-5
0
5
10
15
20
25
30
35
-0.1
0.0
0.1
Angle of Attack (deg)
0.2
0.3
0.4
CL
Figure 19. Euler plus Friction Aerodynamic Performance Predictions at Mach 7 and 115 kft.
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0.5
0.6
0.7
Mach = 7, Altitude = 115kft, AoA = 25º, beta = 0, delta = -5 º
Cart3D
Euler
Overflow Laminar
Navier-Stokes
Figure 20. OVERFLOW Navier-Stokes Solutions Generated to Verify the Validity of Combining Euler CFD
and Friction Drag Predictions.
Mach 7 Lift Curves @ 115kft
Mach 7 Drag Polars @ 115kft
0.7000
0.4500
0.6000
Euler
0.4000
Euler + Flat Plate
0.5000
N-S (Laminar)
0.3500
N-S (Laminar)
0.3000
CD
CL
0.4000
0.3000
0.2500
0.2000
0.2000
0.1500
0.1000
0.1000
0.0000
0.0500
-0.1000
-10.00
0.00
10.00
20.00
0.0000
-0.10
30.00
0.00
0.10
0.20
0.30
0.40
0.50
0.60
CL
Alpha
Mach 7 Pitching Moments @ 115kft
0.0200
0.0150
Euler
0.0100
N-S (Laminar)
Cm
0.0050
0.0000
-0.0050
-0.0100
-0.0150
-0.0200
-10.00
0.00
10.00
20.00
30.00
Alpha
Figure 21. Oveflow Navier-Stokes Solutions Compare Well with Cart3D Euler plus Friction Drag.
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0.70
IX. HIFiRE 4 Trajectory Simulation
The objective of the HIFiRE 4 atmospheric entry trajectory simulation was to start from a provided boost
trajectory at the approximate atmospheric boundary, perform a pull-up maneuver to horizontal flight, then glide to a
landing. This section addresses only the entry phase of the trajectory.
Trajectory analysis and optimization were accomplished using OTIS (Optimal Trajectories by Implicit
Simulation) using its 3-DOF optimization mode. The entry trajectory was analyzed using the 1976 US Standard
Atmosphere model.
Initial conditions for atmospheric entry were selected from the boost trajectory, as follows:
• Altitude = 287,992 ft (87,780 m)
• Velocity = 7115.7 fps (2168.9 m/s)
• Flight path angle = -70.23 deg
Final conditions for the trajectory were at ground impact (landing), as follows:
• Altitude = 0
• Rate of descent = 15 fps (4.572 m/s) or less
• Minimum final speed possible (minimized by OTIS)
Constraints for entry include limiting the dynamic pressure to 20 psf (958 n/m2) or greater, plus keeping the
trajectory within the bounds of the aerodynamic database. An exception to this was allowing the low-speed (i.e.,
below Mach 5) angle-of-attack to extrapolate approximately 2 deg below the zero lower bound of the aerodynamic
database. Trim surface deflections were limited to plus or minus 10 deg for pitch trimming, to maintain a reserve for
control margin.
Figure 22-Figure 26 show time history plots for several parameters of the entry trajectory, which cover 1147
seconds and a total range of 204 nm (378 km). Roll and yaw angles are kept at zero for the entire trajectory analysis.
HIFiRE-4
Re-entry and Pull-up to GAM=0, Then Coast to a Landing
100000
Config: HIFiRE-4
Aero: HIFiRE-4 Aero v7, 5/11/09
Weight= 74.8 kg
Xcg=1.1652 m, Zcg=-0.0696 m
Trimmed
90000
80000
Altitude (m)
70000
60000
50000
40000
30000
20000
10000
0
0
200
400
600
800
1000
Time (sec)
Figure 22. Altitude Time History for the Re-entry Trajectory.
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1200
1400
HIFiRE-4
Re-entry and Pull-up to GAM=0, Then Coast to a Landing
2500
Config: HIFiRE-4
Aero: HIFiRE-4 Aero v7, 5/11/09
Weight= 74.8 kg
Xcg=1.1652 m, Zcg=-0.0696 m
Trimmed
Velocity (m/s)
2000
1500
1000
500
0
0
200
400
600
800
1000
1200
1400
Time (sec)
Figure 23. Velocity Time History for the Re-entry Trajectory.
HIFiRE-4
Re-entry and Pull-up to GAM=0, Then Coast to a Landing
45000
Config: HIFiRE-4
Aero: HIFiRE-4 Aero v7, 5/11/09
Weight= 74.8 kg
Xcg=1.1652 m, Zcg=-0.0696 m
Trimmed
Dynamic pressure (n/m^2)
40000
35000
30000
25000
20000
15000
10000
5000
0
0
200
400
600
800
1000
Time (sec)
Figure 24. Dynamic Pressure Time History for the Re-entry Trajectory.
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1200
1400
HIFiRE-4
Re-entry and Pull-up to GAM=0, Then Coast to a Landing
30
Config: HIFiRE-4
Aero: HIFiRE-4 Aero v7, 5/11/09
Weight= 74.8 kg
Xcg=1.1652 m, Zcg=-0.0696 m
Trimmed
Angle of attack (deg)
25
20
15
10
5
0
-5
0
200
400
600
800
1000
1200
1400
Time (sec)
Figure 25. Angle of Attack Time History for the Re-entry Trajectory.
HIFiRE-4
Re-entry and Pull-up to GAM=0, Then Coast to a Landing
15.0
Trim surface deflection (deg)
10.0
5.0
0.0
-5.0
Config: HIFiRE-4
Aero: HIFiRE-4 Aero v7, 5/11/09
Weight= 74.8 kg
Xcg=1.1652 m, Zcg=-0.0696 m
Trimmed
-10.0
-15.0
0
200
400
600
800
1000
Time (sec)
Figure 26. Trim Surface Deflection Time History for the Re-entry Trajectory.
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1200
1400
IX. Thermal and Structural Analysis
Aeroheating predictions were generated for HIFiRE 4 using the engineering analysis code MINIVER. The
MINIVER code and CHAP, a Boeing proprietary ablation analysis code, are integrated in an automated
aeroheating/TPS analysis and sizing module within Phoenix Integration’s Model Center environment to perform
thermal protection system (TPS) sizing. Aeroheating analysis uses inputs of vehicle geometry and flight trajectory
that includes time histories of Mach number, altitude, angle of attack, and control surface deflection schedule.
Neither CFD solutions nor aerothermal wind tunnel data were used due to limited budget and schedule constraints.
The MINIVER approach is deemed adequate for the short flight duration and all-metallic airframe design of HIFiRE
4. Thermal design is driven primarily by the material thermal capacitance of the airframe. Aeroheating results were
mapped to the 3-D CAD model and subsequent thermal analysis was performed using SINDA. Nonlinear structural
analysis with temperature-dependent material properties was completed by mapping temperatures from the thermal
model using a process developed by Boeing as part of a multi-discipline tool integration effort.
Inboard Surface
Outboard Surface
Btu/ft2-s W/cm2
20.0 22.7
18.0 20.4
16.0 18.2
14.0 15.9
12.0 13.6
10.0 11.4
8.0
9.1
6.0
6.8
4.0
4.5
2.0
2.3
0.0
0.0
Figure 27. Tail Fin Surface Heating.
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Figure 28. Temperature Distribution on Thermal Model.
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Figure 27 shows tail fin surface heating at 27 seconds and temperature distribution at 34 seconds after entry.
The flight encounters peak aerodynamic heating at 27 seconds after entry, while peak airframe temperature lags by 7
seconds, peaking at 34 seconds. Figure 288 shows the temperature distribution on the thermal model. It is interesting
to note that the temperature range is moderate, i.e., less than 500º F, during atmospheric entry. Although the design
of a moderate temperature thermal protection system (TPS) was evaluated, showing that a thin coating of Dyna
Therm DE-350 (Sparesyl) or phenolic cork is effective, TPS was deemed unnecessary for the flight test.
A finite element structural model of HIFiRE 4 was generated with quadratic tetrahedral elements for all
components including the wing, fuselage/body, vertical fins, leading edges, and control surfaces, as illustrated in
Figure 29. The model contains 647,000 nodes and 330,000 elements. Most of the vehicle components will be made
of aluminum Al 6061-T6, although copper C11000 was used for the nose tip and leading edges of wing and body
where higher 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.
Figure 29. HIFiRE 4 Finite Element Model.
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The flight aerodynamic load case used was for finite element analysis was Mach 5.67, 25° angle of attack,
and elevons deflected -6.25°. The temperatures over the flight trajectory were predicted by transient thermal analysis
using SINDA. Temperatures at 34 sec after entry, considered as the maximum thermal load condition, were mapped
to the structural model for subsequent strength analysis. The mapping between the thermal and structural models
was completed utilizing a consistent finite element approach. Based on the results of the thermal analysis, high
temperatures can be observed at the nose tip and on aft areas next to the interface between the fuselage body and
control surfaces, Figure 30.
Figure 30. High Temperature Regions on HIFiRE 4.
Linear strength analysis was performed by combining maximum aerodynamic pressure and thermal loads to
create a worst-case scenario. Temperature dependent material properties for both aluminum and copper were
included in the analysis. The maximum upward deflection predicted at the nose tip is 17.5 mm (0.689 in) and that at
trailing edge is 4.4 mm (0.173 in). Deformation fringe plots and vertical displacements along the centerline of the
bottom surface are illustrated in Figure 31. Von Mises stress distributions for both upper and lower surfaces of the
vehicle are shown in Figure 32, where they are seen in general to be low, i.e., less than 155 MPa (22.5 ksi).
Displacement and stress due to pressure loads alone (i.e., without thermal loads included) were found to be minimal.
A margin check of vehicle strength under loading was conducted using allowables for Al6061-T6 and
C11000, accounting for the temperature dependency on allowable stress. A few very small areas with a negative
margin were observed, which are the result of high temperature and/or high thermal gradients that reduce allowable
stress and increase actual stress in these areas. It is believed that stress relief will prevent these negative margins in
actuality, if real. Moreover, we believe that an analysis with more refined boundary conditions and modeling would
not predict the observed negative margins.
Leading edge deforms upward
by 0.689”, and tail deforms
upward by about .14”
Figure 31. HIFiRE 4 Deflections Under Loading.
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Figure 32. Stress Distribution on Lower Surface of Structure at 34 sec After Entry.
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X. Guidance Navigation and Control
Figure 3 shows the HIFiRE 4 mission profile. At event number 10 the flyers begin running guidance,
navigation and control (GN&C) autocode that commands the surfaces from existing positions to trim at limited rate.
At event number 13 the flyers begin aero closed-loop control upon receipt of a DSTO discreet command.
The initial portion of the trajectory is designed to accomplish a pull-up from a flight path angle of 70 degrees
down to level flight. This requires flying at an angle of attack of 25 degrees as the vehicle enters the atmosphere and
then holding this angle of attack as dynamic pressure builds up and the flight path angle begins to rise. As the
vehicle approaches horizontal flight, angle of attack is reduced to prevent climbing back to higher altitude.
The method described above to achieve the pull-up was modified to avoid an unacceptable level of pitch axis
instability that occurs at high dynamic pressure and low angle of attack. If the vehicle is pitched to lower angle of
attack before it slows sufficiently, it enters an area of the envelope where stabilization by the flight control system is
problematic. On the other hand, if the reduction in angle of attack is delayed until dynamic pressure has decreased to
an acceptable level, the extra lift produced results in the vehicle climbing to unacceptable altitude. To solve this
problem, the pitch down to lower angle of attack is done in two steps: the first is from 25 degrees to 10 degrees, and
then as dynamic pressure drops, the pitch down is continued to -5 degrees angle of attack. In addition to this, the
vehicle is simultaneously rolled to help prevent a climb to higher altitude. Figure 33 shows the area of instability as
well as the pitch and roll schedule used to avoid the instability and prevent a climb to higher altitude. Figure 34 and
Figure 35 show the trajectory with the angle of attack step at 10 degrees along with the original OTIS trajectory.
Following the pull-up to level flight, the vehicle is flown to maintain a dynamic pressure of approximately 75
psf, which results in adequate stability as well as controllability. This portion of the descent trajectory requires some
flight at an angle of attack greater than six degrees for the purpose of controlling dynamic pressure. This is done by
climbing to reduce dynamic pressure and descending to increase dynamic pressure, requiring flight above six
degrees angle of attack. Directional stability for flight above six degrees angle of attack requires the dorsal fin
mentioned previously.
The initial GN&C work was done with an aerodynamic database that did not include the dorsal fin. In that
configuration, it was not possible to fly above six degrees angle of attack at lower Mach numbers during descent
following the pull-up due to yaw axis instability and lack of a rudder. In addition, it was not possible to bank since
this would exacerbate the problem of increasing dynamic pressure due to less lift in the vertical direction. A vertical
fin that stabilizes the yaw axis would therefore enable flight above 6 degrees angle of attack, provide for adequate
control of dynamic pressure by better control of sink rate, limit pitch axis instability to levels which can be stabilized
with current proposed hardware, and allow for banking maneuvers.
The GN&C analysis exposed the need for additional lateral directional stability, which is why the dorsal fin
was added. The dorsal configuration was chosen to conform to the limited space available in the booster fairing.
The aero database with the dorsal fin was discussed in section VIII. We plan to reexamine the GN&C situation with
the new aero database in the future.
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American Institute of Aeronautics and Astronautics
25 Deg
Angle of
Attack
Pitch Down
to 10 Degrees
angle of Attack
Starting at 700 psf
High Drag
High Lift
15 Deg
10 Deg
5 Deg
Pitch Down
to -5 degrees
at 200 psf
70 deg roll
and back to
wings level
0 Deg
High Instability
-5 Deg
0
500
Dynamic Pressure
1000
Figure 33. Pitch and roll schedule used to avoid instability region.
Trajectory (Blue) with Angle of Attack Step at 10 Degrees
And 70 Degree Bank Angle (Red - OTIS)
Dynamic Pressure
Angle of Attack
Dynamic Pressure in Lbs/ft 2 (OTIS Red)
Angle of Attack in Degrees (OTIS Red)
25
1000
900
20
800
15
700
600
10
500
5
400
300
0
200
-5
100
0
-10
0
10
20
30
40
50
60
70
80
90
100
Time in Seconds
0
20
40
60
80
Time in Seconds
Figure 34. Dynamic pressure and angle of attack versus time.
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American Institute of Aeronautics and Astronautics
100
Trajectory (Blue) with Angle of Attack Step at 10 Degrees
And 70 Degree Bank Angle (Red - OTIS)
Bank Angle
Altitude
5
Roll Angle
Altitude in Feet Blue (OTIS Red)
x 10
80
70
2.5
60
50
2
40
1.5
30
1
20
10
0.5
0
0
0
50
100
150
200
250
300
-10
0
50
100
150
Time in Seconds
200
250
300
Figure 35. Altitude and bank angle versus time.
XI.
Conclusions
The Hypersonic International Flight Research and Experimentation (HIFiRE) effort is a joint program
between the United States Air Force Research Laboratory (AFRL) and the Australian Defence Science and
Technology Organisation (DSTO), established through a program agreement. The Boeing Company, the University
of Queensland, the Queensland State Government, BAE Systems Australia, and DLR in Germany are all
collaborators on flights 4, 7 and 8. The objective of HIFiRE is to gather fundamental scientific and engineering data
on the physics and technologies critical to future operational hypersonic flight in several flight experiments using
relatively low cost sounding rockets. The specific objective of HIFiRE 4, the subject of this paper, is to gather flight
data on the aerodynamics, stability, and control of an advanced waverider configuration. HIFiRE 4 is a precursor to
HIFiRE 8, which will add an advanced 3D scramjet engine to the airframe.
The HIFiRE 4 flight experiment will demonstrate strategies for the exo-atmospheric separation and attitude
control, plus endo-atmospheric attitude control of a hypersonic boost-glide vehicle. The experiment will advance
science in the performance of a controlled high angle-of-attack atmospheric entry maneuver using a high
aerodynamic performance waverider configuration. The mission will include the launch of two flyers, one for DSTO
and one for Boeing. Both flyers will conduct the primary experiment of pulling up from a -70 degree flight path
angle to a -50 degree flight path. The Boeing flyer will then continue pulling up to horizontal flight and “land” in the
Woomera dessert. A GN&C strategy has been devised to counter instabilities that have arisen in the configuration.
In future work, the addition of a stabilizing dorsal fin will be more fully evaluated and likely implemented.
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