AIAA 2000-4011
Hyper-X Research Vehicle (HXRV)
Experimental Aerodynamics Test
Program Overview
Scott D. Holland,
William C. Woods,
Walter C. Engelund
NASA Langley Research Center
Hampton, VA
AIAA 18th Applied Aeronautics Conference
August 14-17, 2000
Denver, Colorado
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HYPER-X RESEARCH VEHICLE (HXRV) EXPERIMENTAL AERODYNAMICS
TEST PROGRAM OVERVIEW
Scott D. Holland*, William C. Woods**, and Walter C. Engelund***
NASA Langley Research Center, Hampton, VA
Abstract
This paper provides an overview of the experimental
aerodynamics test program to ensure mission success
for the autonomous flight of the Hyper-X Research
Vehicle (HXRV). The HXRV is a 12-ft long, 2700 Ib
lifting body technology demonstrator designed to
flight demonstrate for the first time a fully airframe
integrated scramjet propulsion system. Three flights
are currently planned, two at Mach 7 and one at Math
10, beginning in the fall of 2000. The research
vehicles will be boosted to the prescribed scramjet
engine test point where they will separate from the
booster, stabilize, and initiate engine test. Following
5+ seconds of powered flight and 15 seconds of cowl-
open tares, the cowl will close and the vehicle will fly
a controlled deceleration trajectory which includes
numerous control doublets tor in-flight aerodynamic
parameter identification. This paper reviews the
preflight testing activities, wind tunnel models, test
rationale, risk reduction activities, and sample results
from wind tunnel tests supporting the flight trajectory
of the HXRV from hypersonic engine test point
through subsonic tlight termination.
Nomenclature
t_ angle-of-attack (degrees)
13 sideslip angle (degrees)
br_f Hyper-X vehicle reference span
C D Drag tbrce coefficient ( Dra_ )
" q_oSref
*Assistant Branch Head, Aerothermodynamics
Branch, Senior Member AIAA.
**Aerothermodynamics Branch, Associate Fellow,
AIAA.
*'*Vehicle Analysis Branch, Senior Member, AIAA.
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CL Lift force coefficient ( Lift )
q_Sref
Cu+,_Lift coefficient derivative with respect to angle
of attack (per degree)
Ci Rolling moment coefficient rolling nzoment )
qooSrt!fbref
Cma Rolling moment coefficient derivative with
respect to aileron deflection (per degree)
Ct!3 Rolling moment coefficient derivative with
respect to sideslip angle (per degree)
Cm Pitching moment coefficient ( pitching moment )
q_Sreflref
Cm,_Pitching moment coefficient derivative with
respect to angle of attack (per degree)
C, Yawing moment coefficient ( yawing moment )
qooSrefbref
C._ Yawing moment coefficient derivative with
respect to sideslip angle (per degree)
C._ Yawing moment coefficient derivative with
respect to aileron deflection (per degree)
Cv Side force coefficient ( side force )
q,,oSref
Cvl3 Side force coefficient derivative with respect to
sideslip angle
Cv_aSide force coefficient derivative with respect to
aileron deflection (per degree)
_rw right lull-flying wing deflection, degrees
_l,_ left full-flying wing deflection, dcgrees
_ aileron deflection (differential horizontal tail:
6,w- 61_L), degrees
_ elevon deflection (symmetric horizontal tail:
(_ru, + 61w)/2 ), degrees
5,, right rudder deflection, degrees
51r left rudder deflection, degrees
_ rudder deflection (6rr +61r)/2, degrees
Ir_f Hyper-X vehicle reference length
qoo freestream dynamic pressure ( l/2p_V_ 2 )
S_f Hyper-X vehicle reference area
AIAA-2(X)0-4011 1
Introduction
The goal of the Hyper-X Program is to demonstrate
and validate the technologies, the experimental
techniques, and computational methods and tools for
design and performance predictions of hypersonic
aircraft with airframe-integrated hydrogen fueled,
dual-mode combustion scramjet propulsion systems
(Ref. 1). Accomplishing this goal requires flight
demonstration of a hydrogen-fueled scramjet powered
hypersonic aircraft. This first-of-its-kind effort is
truly pioneering in that, although hypersonic
propulsion systems have been studied in the
laboratory environment for over 40 years, one has
never before been flight tested on a complete
airframe-integrated vehicle configuration. In order to
meet budget and schedule, thc flight test vehicle
design leveraged existing databases and off-the-shelf
subsystem components wherever possible (Ref. 2).
The design evolution of the Hyper-X configuration
used, as a starting point, the extensive National
Aerospace Plane (NASP) database and experience, as
well as follow-on mission study programs (e.g., Ref.
3). In a sense, the Hyper-X design development was
the reverse of the NASP development. The NASP
program failed to produce a flight vehicle due in part
to insufficient technology development. The Hyper-
X dcsign development looked forward to a 200-foot
operational "'vision vehicle" (developed in the study
of Ref. 4) but sought to design, build, and fly a
minimum size flight research vehicle (as size is a
maior cost driver - Ref. 2) to demonstrate the
technologies and design methodologies necessary to
develop an operational "global reach"
endoatmospheric hypersonic cruise vehicle. Such a
vision vehicle could contribute to key national
civilian and military requirements of routine, cost-
effective access to space, and endoatmospheric,
rapid-response, global reach operations. Preliminary
design studies performed by NASA in early FY95
indicated that a 12 foot vehicle could be "smart
scaled" from the 200 loot operational concept and
still demonstrate scramiet powered acceleration (Ref.
2). Conceptual design trade studies were performed
by McDonnell Douglas Aerospace (MDA - now
Boeing-St. Louis) under contract to NASA (Ref. 5)
between February and May, 1995. MDA completed a
preliminary design between March and October, 1996
under Phase III of the Dual-Fuel Airbreathing
Hypersonic Vehicle Design Study contract (Ref. 6).
This preliminary design, which included basic
structural design, thermal protection system selection,
identification of major system/subsystem components
and potential vendors, preliminary packaging, power
requirements, stage separation approach, booster
integration, and flight test planning, became the
government candidate vehicle for the Hyper-X
program. In July 1996, the Hyper-X program was
approved by NASA Headquarters Code R
(Aeronautics), and a request for proposals (RFP),
based on the government candidate vehicle, was
released in October 1996. The Hyper-X Launch
Vehicle (HXLV) booster development contract was
awarded to Orbital Sciences Corporation in February
1997 and the Hyper-X Research Vehicle (HXRV)
development contract was awarded to MieroCraft,
Inc. in March 1997.
Prior to the release of the RFP, the experimental
aerodynamics program focused on configuration
screening and preliminary database development in
support of control law development and preliminary
trajectory evaluations (including some Monte Carlo
analyses) for inclusion in the RFP. Following
contract award, the experimental aerodynamics
program focused on configuration optimization/
maturation and benchmarking for each phase of the
flight trajectory. This paper will describe the nominal
trajectory and will review the extensive wind tunnel
test program supporting the aero database
development (described in Ref. 7) along that
trajectory.
Mission Profile
The nominal Hyper-X flight trajectories each begin
with a boost to the scramjet engine test conditions on
a modified version of an Orbital Sciences
Corporation Pegasus Hybrid rocket. The HXRV is
attached to the first stage of the Pegasus rocket by
means of a conically shaped adapter. This mated
configuration (thc HXRV, the adapter, and the
booster) is referred to as the Hyper-X Launch Vehicle
(HXLV) or "stack" configuration and is shown in
Figure I.
Figure I. Hyper-X Launch Vehicle (HXLV)
Configuration
The HXLV is carried aloft under the wing of NASA's
B-52 where, in the case of the first two Math 7
AIAA-2000-4011 2
100,000
Altitude, ft
Hyper-X
free flight
Descent
Scram jet
engine start
Hyper-X booster
separation
Booster burn-out
Maneuvers
20,000
Air launch
(over water)
......................... Flight
Termination
Distance
Figure 2. Nominal Mach 7 Hyper-X Flight Profile Superimposed with
Wind Tunnel Test Photographs
experiments, it is dropped at an altitude of
approximately 20.000 ft and a Mach number of 0.5.
Shortly after drop, the booster solid rocket motor is
ignited and the HXLV flies a nominal ascent profile
to the HXRV test point as indicated in Figure 2. At a
point just prior to the scramjet engine test, the Hyper-
X flight vehicle is separated from the launch vehicle.
The entire stage separation sequence, which occurs
over a period of less than 500 milliseconds, presents
several extreme technical challenges in addition to the
basic ones associated with demonstrating the Hyper-
X scramjet engine operation and performance.
Details regarding the stage separation strategies and
associated hardware simulation and testing can be
found in Ref. 8. Details of the experimental test
program for stage separation can be found in Ref. 9.
Immediately following the stage separation event, the
HXRV control system will stabilize the vehicle and
the scramjet test portion of the experiment will begin.
The scramjet engine inlet door will be opened, and
the scramjet fueling sequence will commence. A
combination of silane (Sill4) and gaseous hydrogen
(H2) is injected into the combustor region, resulting in
powered scramjet engine operation. Silane is used
only during the initial ignition process, alter which
pure hydrogen is injected and combusted. Alter the
fuel is depleted, the flight vehicle will record several
seconds of engine-off aerodynamic tare data, then the
inlet cowl door will be shut and the vehicle will
perform a series of aerodynamic parameter
identification maneuvers at hypersonic and
supersonic flight conditions. These maneuvers will
allow the basic aerodynamic stability and control
characteristics of the airframe to be estimated from
the flight data, which will then be compared with the
preflight predictions developed using the ground
based wind tunnel testing and analytical and
computational methods. The vehicle will then fly a
controlled deceleration trajectory, dissipating energy
by performing a series of S-turns. prior to flight
termination at low subsonic conditions.
AIAA-2000-401 I 3