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AMMJk 



A//5 



Al AA 2000-0900 
Aerodynamic Characteristics, 
Database Development and 
Flight Simulation of the X-34 Vehicle 

Bandu NPamadi, 
Gregory J. Brauckmann 
NASA Langley Research Center 
Hampton, VA 

Michael J. Ruth 
Henri D. Fuhrmann 
Orbital 
Dulles, VA 



38th Aerospace Sciences 
IVIeeting & Exhibit 

10-13 January 2000 / Reno, NV 



For permission to copy or republish, contact the American Institue of Aeronautics and Astronautics 
1801 Alexander Bell Drive, Suite 500, Reston, VA 20191 



AIAA-2000-0900 



AERODYNAMIC CHARACTERISTICS, DATABASE DEVELOPMENT 
AND FLIGHT SIMULATION OF THE X-34 VEHICLE 

Bandu N.Pamadi* and Gregory J. Brauckmann/ 
NASA Langley Research Center, Hampton, VA 23681 

Michael Ruth^ and Henri Fuhrmann^ 
Orbital , Dulles, VA 20166 



ABSTRACT 

An overview of the aerodynamic characteristics, 
development of the preflight aerodynamic database and 
flight simulation of the NASA/Orbital X-34 vehicle is 
presented in this paper. To develop the aerodynamic da- 
tabase, wind tunnel tests from subsonic to hypersonic 
Mach numbers including ground effect tests at low sub- 
sonic speeds were conducted in various facilities at the 
NASA Langley Research Center. Where wind tunnel test 
data was not available, engineering level analysis is used 
to fill the gaps in the database. Using this aerodynamic 
data, simulations have been performed for typical de- 
sign reference missions of the X-34 vehicle. 



b 
C. 



h 



NQMFNCLATURE 

Wing span 

Generalized aerodynamic coefficient 

Drag coefficient 

Lift coefficient 

Rolling-moment coefficient 

Pitching-moment coefficient 

Yawing-moment coefficient 

Side-force coefficient 

Height of the moment reference point above the 

ground plane, ft 



* Aerospace Engineer, Vehicle Analysis Branch, 

Aerospace Systems Concepts and Analysis 

Competancy, Associate Fellow AIAA. 

'Aerospace Engineer, Aerothermodynamics Branch, 

Aerodynamics and Aerothermodynamics Competancy, 

Senior Member, AIAA. 

^Technical Manager for Guidance, Navigation and 

Control, Member AIAA. 

^Aerospace Engineer, Aerodynamics, Member AIAA, 

Copyright © 2000 American Institute of Aeronautics and 
Astronautics, Inc. No copyright is asserted in the United States 
under Title 17,U.S.Code. TheU.S. Govemmenthasaroyalty- 
free license to exercise all rights under the copyright claimed 
herein for Governmental purposes. All other rights are reserved 
by the copyright owner. 



M 
a 

5 

e 

I' 

r 
sh 

AC 



Mach number 
Angle of attack, deg 
Angle of sideslip, deg 
Aileron deflection angle, deg 
Elevon deflection angle, deg 
Body flap deflection angle, deg 
Rudder deflection angle, deg 
Speedbrake deflection angle, deg 
Incremental in generalized aerodynamic 
coefficient C. 

INTRODUCTION 



The X-34 vehicle being developed by the Orbital 
Sciences Corporation for National Aeronautics and Space 
Administration (NASA) is an integral part of the reus- 
able launch vehicle (RLV) technology program current- 
ly being pursued by NASA with industry partnership. A 
schematic representation of the RLV technology dem- 
onstration path is shown in Figure 1 . The primary goal 
of the RLV program [I] is to develop key technologies 
that will significantly lower the cost of access to space. 
The X-34 program originally started in spring of 1995 
when the team of Orbital Sciences Corporation and Rock- 
well International was awarded a NASA contract to build 
an unmanned, fully reusable, two-stage, orbital vehicle 
capable of delivering approximately 1500 lb payload to 



RLV 

2,000,000 fb 
Earth Orbit 



X-34 

45,000 lb 
Mach 8 



X-33 

273,000 [b 
Mach 15 




DC-XA 

42,000 lb 
Subsonic 



Figure L RLV technology demonstration path. 



I 



American Institute of Aeronautics and Astronautics 



low-earth orbit. However, the program was cancelled in 
late 1 995 when Orbital Sciences Corporation and Rock- 
well determined that the program was not economically 
feasible. This program was resurrected in spring of 1 996 
when NASA solicited proposals on a different vehicle, 
also designated X-34 [2]. Orbital Sciences Corporation 
(now Orbital) was awarded this contract in June 1996. 

The current X-34 vehicle is an unmanned suborbit- 
al, technology demonstrator vehicle capable of reaching 
an altitude of 250,000 ft and a speed of Mach 8, Some of 
the key technologies related to RLV that will be demon- 
strated by the X-34 vehicle include primary and second- 
ary composite structures, advanced thermal protection 
systems (TPS), low cost avionics, rapid turn around 
times, autonomous flight including landing, and all 
weather airplane-like operations. 

The NASA Langley Research Center (LaRC) is in- 
volved in the aerodynamic analysis, wind tunnel testing 
from subsonic to hypersonic speeds and the development 
of the preflight aerodynamic database of the X-34 vehi- 
cle. Orbital is responsible for the flight simulation of the 
X-34 vehicle. An analysis of the X-34 wind tunnel test 
data up to Mach 6 was reported in [3]. The formulation 
and development of the aerodynamic database was dis- 
cussed in [4]. Since then, Mach 10 wind tunnel tests have 
been performed and with this, all the planned wind tun- 
nel tests have been completed and the database has been 
updated. Orbital has performed numerous simulations 
for various design reference mission (DRM) trajecto- 
ries that are expected to be flown during the X-34 flight 
test program. The objective of this paper is to present an 



overview of these activities and discuss salient aerody- 
namic and flight characteristics of the X-34 vehicle. 

VEHICLE/MISSION DESCRIPTION 

The X-34 vehicle has a close similarity with the 
Space Shuttle Orbiter but is relatively smaller in size. A 
schematic three-view diagram of the X-34 vehicle is pre- 
sented in Figure 2. The X-34 vehicle has an overall length 
of about 58 ft, wing span of 28 ft and a height of about 
12 ft. The approximate gross weight of the X-34 vehicle 
is 45000 lb. The main 'wing of the X-34 vehicle has a 
leading edge sweepback of 45°, a dihedral of 6°, and an 
80° leading edge strake and full span split elevons (from 
actuator torque considerations). The elevons on the same 
side are always deflected together. Deflected symmetri- 
cally, elevons produce pitch control and asymmetric 
deflections provide roll control. The vehicle has a body 
flap located at the trailing edge of the fuselage. The body 
flap helps to shield the engine nozzle from aerodynamic 
heating at hypersonic speeds and also augments pitch 
control. The vehicle features a centerline, all movable 
vertical tail for directional stability/control. The vehicle 
also features reaction control system (RCS) jets located 
at the aft end of the fuselage for roll and yaw control 
when the vertical tail becomes ineffective at high alti- 
tude and high Mach number (low dynamic pressure and 
high angles of attack) flight conditions. The vertical tail 
has a split speedbrake like the Space Shuttle Orbiter for 
energy management during descent. The TPS system on 
X-34 consists of a mix of ceramic tiles and blankets. 
Ceramic tiles are used in the stagnation regions of the 
nose and wing leading edges where the aerodynamic 




332,9" 



Reference Numbers | 


Sref 

mac 


357.5 ft2 


174.48 in. 


bref 


332 in. 


mrc 


420 in. from the nose 





^ 12,r Static Taif Scrape 
i -J- Static Ground Une 



2^^ Static 
f Incidenoe 



JL 



>''f 



1 


t 

\ 

142.2' 

i 


&s.^ 


- 1 


1 *w 





JL 



Figure 2. Schematic diagram of the X-34 vehicle. 



American Institute of Aeronautics and Astronautics 



heating is quite severe. Tiiree types of blankets are em- 
ployed for the rest of the acreage of the vehicle depend- 
ing on the anticipated thermal environment. Additional 
information on the TPS can be found in [5]. 

The X-34 will be powered by the "Fastrac" rocket 
engine which is under development at the NASA Mar- 
shall Space Flight Center (MSFC), Huntsville, Alabama. 
The bi-propellent (liquid oxygen (LOX) and RP (kero- 
sene)) Fastrac engine is designed for a nominal thrust of 
60,000 lb and is expected to have a thrust vectoring ca- 
pability of ±1 5"^ in the pitch plane. The X-34 vehicle has 
one composite RP tank and two aluminum LOX tanks 
located axially one behind the other. The RP tank is lo- 
cated in the front part of the fuselage and ahead of the 
two LOX tanks. 



Early in 2000, the X-34 vehicle will undergo run- 
way tow testing at the DFRC. These tests will be used to 
prove out the^autonomous landing, steering, and brak- 
ing algorithms. The unpowered X-34 vehicle will be 
towed up to a speed of about 80 mph with a truck and 
released. Tracking of the runway centerline, steering and 
braking effectiveness will all be monitored during the 
runway tow test. Following successful completion of the 
runway tow test, some additional L-101 l/X-34 captive 
carry flights and "dry run" releases will be performed to 
prepare for the first approach and landing test in the 
spring of 2000. These approach and landing tests will be 
conducted at the White Sands Space Harbor in New 
Mexico and will validate the release, approach, landing, 
and rollout phases of the X-34 flight profile. 



A typical X-34 mission profile is depicted in 
Figure 3. The X-34 vehicle will be "captive" carried 
under the belly of the L- 1 01 1 aircraft up to an altitude of 
about 38,000 ft and a Mach number of 0.7 at which point 
it will be released. The vehicle will be unpowered and 
all its control surfaces will be locked for about one sec- 
ond following the drop. Once the vehicle makes a safe 
separation from the L-IOl 1 aircraft, the Fastrac engine 
will ignite and accelerate the vehicle towards its target 
altitude of 250,000 ft and target speed of Mach 8. After 
engine burn out, the vehicle will coast and glide back to 
earth and execute an autonomous, airplane-type landing 
on a conventional runway. 



Bumout^- 

L-1011 / 

^ carrier aircraft ^ 
Uunch __ yfAscent 



Descent 



Ignition after 
separation maneuver 




Recovery 



Down range landing 



Figures, Typical X'34 flight profile. 

VhlQm TEgT PHOQRAM 

The X-34 flight test program includes L-1011 cap- 
tive carry testing, runway tow testing, unpowered ap- 
proach and landing tests, and incremental powered en- 
velope expansion flights up to the full Mach 8 capabili- 
ty. The captive carry tests serve to validate and provide 
FAA (Federal Aviation Administration) certification of 
the L-1011 as the carrier vehicle for the X-34. These 
tests were conducted at the NASA Dryden Flight Re- 
search Center (DFRC), California in the fall of 1999 up 
to the maximum captive carry flight Mach number of 
0.87. 



The Fastrac engine static fire testing will be con- 
ducted at the DFRC during the spring of 2000 and will 
be followed by low Mach powered flight of the X-34 
vehicle in the summer of 2000. Subsequently, several 
flights will be conducted to gradually expand the flight 
envelope of the X-34 vehicle. It is proposed to collect 
aerodynamic data in these tests and use it to update the 
X-34 aerodynamics database for subsequent flights. 
Emphasis will be placed on envelope expansion, not 
operability, at this point in the program. The flight test- 
ing of the X-34 vehicle will then move to NASA 
Kennedy Space Center (KSC), Florida which will serve 
as the proving ground for X-34 operability. The X-34 
vehicle will perform high Mach number flights off the 
Eastern coast from North Carolina down to the KSC. 
Flights will be conducted every two weeks for a three 
month proving period with minimal ground crew. A 
surge capability will also be demonstrated in which the 
vehicle will be turned around and flown within 24 hours 
which is an important requirement of the X-34 program. 

The X-34 program is currently planning for 27 
flights. These flights include experiments to demonstrate 
new technologies in TPS, structures, and composite liq- 
uid oxygen tanks. Additional flight experiments may 
include thermal and pressure measurements for valida- 
tion of computational fluid dynamics and computation- 
al aerodynamic heating codes. 

WTND TUNNEL TEST FACILITIES 

A brief description of various LaRC wind tunnel 
facilities used in generating the lest data included in the 
X-34 aerodynamic database is presented below. Addi- 
tional information on these LaRC lest facilities may be 
found in [6,7,8,9]. The L-101 1^^-34 captive carry and 
separation aerodynamic model tests were conducted by 



American Institute of Aeronautics and Astronautics 



Orbital in the Calspan transonic wind tunnel facility. 
Also, some tests on the X-34 model were conducted in 
the trisonic wind tunnel facility at MSFC. These test re- 
sults are not discussed in this paper. 

LaRC 14- bv 22-Foot Subsonic Tunnel 

The Langley 14- by 22-Foot Subsonic Tunnel is a 
closed circuit, single return, atmospheric tunnel with a 
maximum speed of 338 ft/sec. The test section measures 
14.5- by 21.8 ft and has a length of about 50 ft. The 
maximum unit Reynolds number is 2,1 x 10^ per ft. The 
tunnel is equipped with boundary layer suction on the 
floor at the entrance to the test section. 

LaRC 16-Foot Transonic Tunnel 

The Langley 16-Foot Transonic Tunnel is a closed 
circuit, single return, continuous flow atmospheric tun- 
nel. The test medium is air. This tunnel has a slotted 
wall, octagonal test section which measures 1 5.5 ft across 
the flats. The normal test Mach number ranges from 0.3 
to 1.3. The angle of attack can be varied up to 25°. The 
unit Reynolds number varies from 2,0 to 4 x 10^ per ft. 

LaRC Unitary Plan Wind Tunnel 

The Langley Unitary Plan Wind Tunnel (UPWT) is 
a continuous flow, variable pressure, closed circuit pres- 
sure tunnel having two separate test sections, called low 
Mach number test section (leg 1) and high Mach num- 
ber test section (leg 2), Each test section measures 4- by 
4 ft and has a length of 7 ft. The tunnel is capable of 
operating from near vacuum conditions to a pressure of 
10 atmospheres. The low Mach number test section cov- 
ers the Mach number range from 1 .46 to 2.86 and the 
high Mach number test section from 2.3 to 4.6, The an- 
gle of attack capability is from -12° to 22° with possi- 
bility for testing at higher values using dogleg strings. 
TTie unit Reynolds numbers range from 1 .0 to 4.0 x 10^. 

LaRC 20-Inch Mach 6 Tunnel 

The Langley 20-Inch Mach 6 Tunnel is a blow down 
test facility that uses heated, dried and filtered air as the 
test medium. The test section measures 20.5- by 20-inch- 
es. Typical operating stagnation pressures range from 
30 to 500 psi and the stagnation temperature from 750° 
to i(X)0° R. The unit Reynolds numbers range from 0.5 
to 8 X 10^ per ft. This tunnel has a capability to run 
continuously up to 15 minutes. The tunnel is equipped 
with a model injection system on the bottom of the test 
section that can insert a sheltered model into the air 
stream in less than 0.5 seconds. 



LaRC 3Mnch Mach 10 Tunnel 

The Langley 31 -Inch Mach 10 Tunnel is a hyper- 
sonic blow down facility that uses dried, heated, and fil- 
tered air as the test gas. The facility typically operates at 
stagnation pressures from 350 to 1450 psia and at a stag- 
nation temperature of I850°R, with corresponding free 
stream unit Reynolds numbers of 0.5 to 2.2 x 10^ per ft. 
A three-dimensional, contoured nozzle is used to pro- 
vide a nominal freestream Mach number of 1 in the 3 1 - 
Inch square test section. A side-mounted model injec- 
tion system can insert models from a sheltered position 
to the tunnel centerline in less than 0.6 sec. Run times 
up to 3 minutes are possible with this facility although 
current test run times were on the order of one minute. 

MODELS. INSTRUME NTATION 
AND TEST PROCEDURE 

The model for the 14-ft by 22-ft low subsonic, free 
stream and ground effect tests was a iO% scale model of 
the X-34 outer mold line (OML) geometry inclusive of 
TPS. The test model had remote activation of elevons, 
body flap and rudder. The floor boundary layer suction 
was used in the X-34 ground effect tests. The ground 
effect test data was obtained for various separation 
heights (measured from moment reference point to the 
ground plane) ranging from 0.3 to 2.5 times wing span. 
The X-34 vehicle has two doors for the main gear, one 
on each side, but a single door for the nose gear, only on 
the left side. Therefore, when the nose gear is down and 
its door is open, the configuration becomes aerodynam- 
ically asymmetric giving rise to side force, rolling and 
yawing moments at zero sideslip. 

The model for the 16-Foot Transonic Tunnel and 
the UPWT was a 0.033-scale model of the X-34 OML 
geometry, for the 20-Inch Mach 6 Tunnel was a 0.018 
scale model of the X-34 OML geometry and that for the 
3 1 -Inch Mach 1 Tunnel was a 0.01 3-scale model of the 
X-34 OML geometry. 

For the test models in the 14-by 22-Fool Subsonic 
Tunnel, the 16-Foot Transonic Tunnel and the UPWT 
(leg 1 ) tests, boundary layer transition trips were applied 
at the nose and the leading edges of the wing and verti- 
cal tail to promote turbulent flow over the test models. 
The models tested in the UPWT (leg 2), the 20-Inch Mach 
6 Tunnel and the 31 -Inch Mach 10 Tunnel were not 
tripped. However, the data from the UPWT (Leg I ) tests, 
where models with and without the trips were tested, 
showed that tripping had little effect on lift and pitching 
moment coefficients but resulted in a drag coefficient 
increase of about 2% for Mach 1 .6 to 2.5. 



American Institute of Aeronautics and Astronautics 



The aerodynamic forces and moments were mea- 
sured using six component strain gage balances. The 
balances used in the 20-Inch Mach 6 Tunnel and 31- 
Inch Mach 10 Tunnel were water cooled to minimize 
the balance temperature variations due to aerodynamic 
heating. Corrections were applied to the balance mea- 
surements to account for the temperature effects only 
for the 3 1 -Inch Mach 1 Tunnel test data. 

The force and moment data were acquired in a "pitch 
and pause" manner. The balance moment reference cen- 
ter, expressed in terms of full scale vehicle, was located 
at 420 inches from the nose. The base and cavity pres- 
sures were measured on all models except the model in 
the 31 -Inch Mach 10 tests and these were used to make 
correction to the measured axial force. In the Mach 10 
tests, owing to limitations of the model and cavity size, 
the cavity pressure could not be measured and no cor- 
rection to the axial force was made. 



FORMULATION OF AERODYNAMIC 
DATABASE 

An important aspect of developing the aerodynam- 
ic database is the formulation of suitable aerodynamic 
models. The development of aerodynamic models for 
the evaluation of the static aerodynamic forces and mo- 
ments of the X-34 vehicle in free flight and for flight in 
ground effect is discussed in the following. This discus- 
sion does not include the control surface hinge moments 
and the dynamic or damping derivatives. This formula- 
tion is similar to that used in the Space Shuttle Orbiter 
data book [10]. 

Aerodynamic Coefficients in Free Flight 

By free flight, it is meant that the vehicle is out of 
ground effect. This assumption generally holds when the 
vehicle is at a height exceeding 2.5 wing spans. 



In general, the tests in all the above facilities cov- 
ered elevon deflections from -30° to -i-20°, aileron de- 
flections of -30° to +20^ ( elevons on one side deflect- 
ed, those on the other side held at zero), body flap de- 
flections of-15° to +20°, rudder deflection of S'^ to 30° 
and nominal speedbrake deflections of 30° to 90°. For 
subsonic and low supersonic tests (up to Mach 2.5), the 
angle of attack varied from -4° to 20°. For UPWT (leg 
2) and Mach 6 tests, the angle of attack reached up to 
36°. However, for Mach 1 tests, the maximum angle of 
attack ranged only up to 28°. The sideslip was in the 
range of -6° to +6° for tests in the 14-by 22-Foot Sub- 
sonic Tunnel, the 16-Foot Transonic Tunnel and the 
UPWT. For Mach 6 tests, the sideslip was in the range - 
3° to +4°. In all the tests up to Mach 6, the lateral/direc- 
tional test data was obtained for angle of attack fixed 
with sideslip variations as well as sideslip fixed with 
angle of attack variations. However, for the Mach 10 
tests, the lateral/directional test data was obtained with 
sideslip fixed at -3° and +3° and angle of attack varying 
from to 28°. 

The uncertainties in the balance measurements for 
various Mach numbers were estimated as follows: nor- 
mal force from 0.00 1 to 0.02 1 6, axial force from 0.0008 
to 0.0054, pitching moment coefficient from 0.004 to 
0.0177, side force coefficient from 0.0028 to 0.0179, 
rolling moment coefficient from 0.0005 to 0.001 1 and 
the yawing moment coefficient from 0.0008 to 0.004. 
Additional information on the measurement uncertain- 
ties can be found in [3]. 



Assume that the vehicle is operating at a combined 
angle of attack and sideslip. Let C. represent any one of 
the six aerodynamic coefficients ^ ^^ lyC ^,C y,C ^ or C^ 
and be given by 



C,,ou„ = C,,(oc, M) + AC. , + AC. , + AC. 



V 



+ AC,, +AC., +AC.,^ + AC.,,^ 



+ AC o fl + AC. <r « + AC , ^ o 

1.8^^ i,S^fj,fi i,LG,l3 



(1) 



Here, C. . , is the total coefficient of the vehicle and is 
expressed as a sum of its value for the baseline at angle 
of attack (zero sideslip) C.^^ia, Af), and various incre- 
mental coefficients due to deflection of control surfaces 
like elevons (5,), ailerons (5^), body flap (S^j), rudder 
(5^), speedbrake (5^^) or the extension of landing gear 
(LG), all in zero sideslip (P = 0) plus the incremental 
coefficients due to sideslip for the baseline, deflection 
of rudder, speedbrake, and extension of the landing gear 
in the presence of sideslip. It is assumed that the sideslip 
has effect only on the baseline, and when the rudder and 
speedbrake are deflected but has no effect when elevons, 
body flap or ailerons are deflected. 

The parameter AC- ^ denotes the incremental coef- 
ficient due to a elevon deflection and is defined as 



AC,^=C.(a,M,5^)-C.,(a,Af) 



(2) 



American Institute of Aeronautics and Astronautics 



The other incremental coefficients due to the deflection 
of body flap (AC, ^ J, rudder (AC^ 5 ), speedbrake (AC, 5^^) 
and landing gear (AC,j;^^) are defined in an identical 
manner as in equation (2). The parameter AC,^ repre- 
sents the incremental coefficient due to aileron deflec- 
tions and is defined in a slightly different manner. For 
lift, drag and pitching moment coefficients 



AC., = 0.5(AC. , ^, + AC,, __, ) - AC,, 

a e e,L e e,R e 



(3) 



Thus, to evaluate AC^^ , the elevon aero data is used 
twice, once assuming 5 = 5,^ to obtain AC^^^^ and 
then assuming d = 5 „ to determine ACij _s „ . As a 
check, when aileron deflection is zero, i.e., o^^ = o^^, 
AC- ^ = as expected. 

The incremental coefficients due to sideslip are eval- 
uated as follows: 

For the baseline in sideslip, the incremental coefficient 
is defined as 



AC.^.= C.{a.M,p)-C.(a,M) 



(4) 



The incremental coefficient due to the deflection of 
rudder when the vehicle is in sideslip is defined as, 

^^i.s^0 = [<^r (« . A^. A 5, ) - C. (a ,M,p)]~ AC. ^^ (5) 

The incremental coefficients due to speedbrake deflec- 
tion or the extension of the landing gear are defined in 
an identical manner as in equation (5). 

The formulation as given by equation (1) is of gen- 
eral nature. Usually, some of the incremental coefficients 
are zero. For example, AC^, ^^ = "^^v.^j^^^ ~ ^^yA,h " ^' 
Additional details on the fonnulation of the free fii 
aerodynamic database may be found in [4]. 

Aerodynamic Co^fTicients In grwn^ Effect 



ght 



Consider the vehicle with its landing gear fully ex- 
tended and operating in the proximity of the ground 
(h^ < 2.5). Here, h is the height of the vehicle above the 
ground plane, assumed equal to the vertical distance 
between the moment reference point and the ground 
plane, and b is the wing span. Let 

C(a, A 5,5,^, 5,, 5^, 5 , h/b) = C.{a,h/b = 00) 

+ AC,.(a , A 5 , 5,^, 5,, 5^, 5 , h/b) (6) 

Here, it is assumed that the aerodynamic coefficient in 
ground effect is expressed as a sum of its value in free 



flight (h/b = 00) for the baseline at angle of attack (zero 
sideslip) and an incremental coefficient due to the de- 
flection of the control surfaces and ground effect. The 
inclusion of the term h/b in the parenthesis denotes that 
the coefficient C, is evaluated in ground effect. As be- 
fore, C denotes any one of the six aerodynamic coeffi- 
cients C,yCr.yC , C ^ , C . oT C . Assumc that the incre- 
mental coefficient in equation (6) is given by 

AC.(a, A 5,, 5,^, 5,, 5^, 5 ,M?) = AC,(a, M?) 

+ AC. (a , 5,, h/b) -h AC (a , 5^^, h/b) 

+ AC. (a , 5^,, hjly) + AC (a , 5^, h/h) 

+ AC. (or , 5^, h/b) + AC. {a, p, h/b) 

+ AC. (a , A 5,. M>) + AC.(a, A 5. hA>) (7) 

Here, AC. (a, h/b) represents the incremental coefficient 
for the baseline at angle of attack and in the presence of 
the ground with respect to the baseline in free flight at 
the same angle of attack and is defined as 



AC.{a,h/b) = C.(a,h/b)-C,(a,h/b = oc) 



(8) 



The parameter AC. (a, 5,, h/b) represents the incremen- 
tal coefficient due to elevon deflection at angle of attack 
and zero sideslip and in the presence of the ground with 
respect to the baseline in zero sideslip at the same val- 
ues of a, h/b and is defined as, 



Aq{a,S^,h/b) = q(a,5^,h^)-q(a,h/b) 



(9) 



The incremental coefficients due to the deflection of body 
flap, speedbrake, ailerons and rudder are defined in an 
identical manner as given in equation (9). Next, consid- 
er the incremental coefficients involving sideslip. The 
incremental coefficients due to sideslip are defined as, 

AC,{a , P,hA>) = C.{a . P,hA))- C.{a .h/b) (10) 

AC(a, A 5r. hA» = [C.(a, A 5 , M>)- C (a, A h/b)] 

-AC(a,5^,M?) (11) 

AC.(a,A5,,,M>) = [C.(a,A5,,M^)-C.(a,A^)] 
-AC(a,5,,M^) (12) 

Additional details on the formulation of ground 
effect aerodynamic model may be found in [4]. 



American Institute of Aeronautics and Astronautics 



Process of Developnieiit of 
the Aerodynamic Database 

The current X-34 program started in the summer of 
1996. At that time, some preliminary wind tunnel test 
data at Mach 0,2 and at Mach 6 were available for the 
previous (cancelled) X-34 configuration. While the wind 
tunnel test program on the new X-34 configuration was 
yet to start, it was necessary to quickly put together an 
aero database for the guidance, navigation and control 
engineers to get started with the flight control system 
design. For this purpose, the first version of the aero 
database was developed using APAS (Aerodynamic Pre- 
liminary Analysis System) which is an interactive com- 
puter code capable of giving quick engineering estimates 
from subsonic to hypersonic speeds [11,12]. The APAS 
predictions were adjusted using the available wind tun- 
nel data at Mach 0,2 and 6,0 for the previous version of 
the X-34. For other Mach numbers, past experience with 
similar vehicles such as the Space Shuttle Orbiter and 
other wing-body configurations was used to anchor the 
APAS predictions [13,14,15]. 

Subsequently, the aero database was regularly up- 
dated by replacing the APAS results with the wind tun- 
nel test data as and when such data on the current X-34 
model became available. All the planned wind tunnel 
tests were completed and the final update to the aerody- 
namic database was accomplished in October 1999. 

As said before, the lateral/directional data from the 
31 -Inch Mach 10 tests were obtained only for sideslip 
of -3° and +3°. In view of this, several gaps exist in the 
Mach 10 test data. To fill these gaps and populate the 
database at Mach 10, APAS was used. The approach 
taken was to run APAS for Mach 6 and Mach 10, calcu- 
late the incremental coefficient due to Mach number 
variation from 6 to 10 when all other parameters remain 
constant. Next, add this incremental to the Mach 6 test 
data so that the Mach 6 test data is made applicable for 
Mach 10. As an example, consider the aerodynamic co- 
efficient for the baseline vehicle at combined angles of 
attack and sideslip, 

C.(a. A Af 10) = C,(a, A A/6, WT) -h AC.(APA^ (13) 

where 



AC. (APAS) = C.{a, A ^'0, APAS) 
-C.{a,p,M6,APAS) 



(14) 



each of the terms appearing in the free fiight and ground 
effect aerodynamic models. For the free flight aero da- 
tabase, the Mach number ranges from 0.3 to 10.0 with 
closely spaced values in the transonic regime. The angle 
of attack varies from -6° to 21° for M = 0.3 to 2.5 and 
from -5° to 40° for M = 3.0 to 1 0.0, The data is present- 
ed for elevon deflections (positive downwards) of -30° 
to 20°, aileron deflections from -30° to -h20° (left elevons 
deflected, right held at zero), body flap deflections of 
-15° to 20°, rudder deflections (positive to left) from 
-5° to -20° and nominal speedbrake deflections from 
30° to 90°. The sideslip ranges from ^° to +5°.The 
ground effect aerodynamic data is presented for Mach 
0,3 and h/b varying from to 2.5. The control deflec- 
tions considered in the ground effect aerodynamic data- 
base are similar to those in the free flight aero database. 
All the aerodynamic data in the database is with respect 
to the moment reference point located at 420 inches from 
the nose. 

RESULTS AND DISCUSSION 

Aerodynamic Characteristics 

Some of the salient aerodynamic characteristics of 
the X-34 vehicle are discussed in this section. For more 
details reference may be made to [3,4]. 

The variation of lift coefficient and pitching mo- 
ment coefficient with angle of attack at various Mach 
numbers are presented in Figures 4 and 5. It is observed 
that the vehicle does not encounter stall up to 21° angle 
of attack in subsonic/supersonic range and up to 40° at 
hypersonic speeds. The vehicle is unstable at low speeds 
(M = 0.3) in pitch at low a, exhibits a pitch up tendency 
around a = 9° and then a stable break with further in- 
crease in a. The vehicle becomes more stable at tran- 




-10 5 



5 10 15 20 25 30 35 40 
Angle of attack, deg 



The aerodynamic data in the aero database is pre- 
sented in the form of tables so that the user can evaluate 



Figure 4. Variation of lift coefficient with angle of 
attack for various Mach numbers. 



American Institute of Aeronautics and Astronautics 



sonic/supersonic speeds and the angle of attack at which 
pitch up occurs also increases as observed in Figure 5. 
At hypersonic speeds, the vehicle becomes unstable be- 
cause of the increasing lift developed by the forward parts 
of the fuselage and exhibits a tendency for a stable break 
at high angles of attack. This type of variation in pitch- 
ing moment coefficient is typical of wing-body config- 
urations at hypersonic speeds. 

The variation of untrimmed lift-to-drag ratio is pre- 
sented in Figure 6. It is observed that at low subsonic 
speeds, the vehicle has a lift-to-drag ratio of as much as 
8 at low angles of attack. However, as Mach number 
increases the value of lift-to-drag ratio decreases and 
assumes values ranging from I to 2. 

An example of elevon effectiveness from subsonic 
to hypersonic speeds is shown in Figure 7 for two val- 
ues of angles of attack, a - 6° and 20''. For a = 6°, it is 
observed that the elevon effect decreases rapidly at su- 



-0.25 



— } 1 1 " T -- 11 1 I 1 1 


M 

.,u 




— * i 




"~'^v^^J^^J_M = 0.3 ] 




^ 


"^S/^^^^^^^ 


-■[ 




r\V : ; . josLL \ I. 




;- 




i 



-10 5 5 10 15 20 25 30 35 40 
Angle of attack, deg 



Figure 5, Variation of pitching moment coefficient 
with angle of attack for various Mach numbers. 



personic and hypersonic speeds. It is interesting to note 
that for a = 20°, the downward deflected elevons still 
retain their effectiveness all the way up to Mach 10. 

The variation of body flap effectiveness for a = 6° 
and 20"^ is shown in Figure 8. The data for the body flap 
deflection of -15° goes only up to Mach 4.6, As ob- 
served above for elevons, the body flap effectiveness 
decreases at supersonic/hypersonic speeds for a=6° and 
for a = 20°, the downward deflected body flap retains 
effectiveness all the way up to Mach 10. 

Typical aileron effectiveness as measured by the 
rolling moment coefficient is shown in Figure 9 for 
a = 6° and 20°, It is observed that for a= 6°, the aileron 
effectiveness decreases at supersonic and hypersonic 
speeds and for a = 20, the downward deflected ailerons 
retain their effectiveness all the way up to Mach 10. 

The rudder effectiveness as measured by the yaw- 
ing moment coefficient is shown in Figure 10 for a = 6° 
and 20°. It is observed that the rudder effectiveness in- 




4 5 6 
Mach number 



Figure 7. Elevon effectiveness at a= 6^ and a= 20^ 



9 
8 

7 
6 

5 

Lift-to-drag . 
ratio 

3 
2 

1 



-1 



, , 



a, deg 
I 6 

: + 9 
; • 15 

; " 21 

: • 25 
• 30 

: • 35 



I t i i 

, ii: ? \ ?.,. 



Q r- 



1 23456789 10 
Mach number 




Figure 6. Variation of lift-to-drag ratio (untrimmed) 
with Mach number for various angles of attack . 



Figure 8, Body flap effectiveness at 
a ^O"" and a =20'', 



8 
American Institute of Aeronautics and Astronautics 




0.04 

0.02 

^C,,5a 

-0.02 

-0.04 

-o.oe 
-o.oa 



r\ ' ; * 


I 1 1 


^.■^^~ 


.--#.- - 


-in. 


^'"Vt**? T"T 






.. ..^.... ^,..;. „..,„, ,^... 


j 


...j„„ 


....; 


or" '- 


=™|=HM- 9 — 

; = 20^^ i 


t 


„.|.... 


\ 



Sg, deg 

• 20 
+ 10 
« -10 

♦ -20 
o -30 



4 5 6 
Mach number 



10 




3 4 5 6 

Mach number 



Figure 9. Aileron effectiveness at 
a = d"" and a = 20°. 



Figure IL Speedbrake effectiveness at 
a ^6° and a ^20°, 




x103 






umm^. j^ 


^^^^^^^ 


#^ 


:stable 




;a = 6°^ 





4 5 6 
Mach number 




P. (leg 
» 1 

* 2 
X 3 
o 4 

• 5 



4 5 6 

Mach number 



Figure 10, Rudder effectiveness at 
a = 6"" and a = 20"", 



Figure 12. Roiling moment coefficient due to sideslip 
for the baseline configuration at a= 6° and a = 18° 



creases at transonic speeds but decreases rapidly at higher 
Mach numbers. At a= 20, the rudder is virtually inef- 
fective above Mach 5. In such situations, the X-34 flight 
vehicle will make use of the RCS for directional control 

The speedbrake effectiveness as measured by the 
drag incremental also varies in a similar fashion as shown 
in Figure 1 1 . The increment in drag due to speedbrake is 
also accompanied by an increase in pitching moment 
which can augment the pitch control. The loss of rudder 
and speedbrake effectiveness at high angles of attack 
and high Mach numbers is due to the immersion of these 
surfaces in the low pressure wake of the fuselage and 
wings. 

The lateral and directional stability characteristics 
for a= 6° and 18° are shown in Figures 12 and 13. It is 
observed that for a = 6°, the vehicle is stable in roll 
iCfQ< 0) up to about Mach 1 .7 and beyond Mach 1 ,7, it 
becomes unstable in roll (Cfg > 0). For a= 18°, the ve- 
hicle is stable in roll at all Mach numbers (except around 
Mach 1.0) due to the increasing stabilizing effect pro- 




4 5 6 

Mach number 



Figure 13. Variation of yawing moment coefficient for 
the baseline configuration at a- 6° and a = 18°, 



vided by the wing dihedral. For a = 6°, the vehicle is 
directionally stable (C « > 0) up to Mach 1 .5 and unsta- 
ble (C^Q< 0) beyond Mach 1.5 as shown in Figure 13. 
For higher angles of attack (a = 18°) the vehicle be- 
comes directionally unstable at all Mach numbers. 



American Institute of Aeronautics and Astronautics 



The effect of landing gear deployment at low sub- 
sonic speeds (M = 0.3) is shown in Figures 1 4 and 1 5. It 
is observed that the landing gear deployment leads to a 
more nose down pitching moment up to 12° angles of 
attack and then the trend reverses at higher angles of 
attack. These incremental coefficients correspond to 
about half a degree of elevon deflection. Further, the ve- 
hicle experiences significant asymmetry in the variation 
of pitching moment coefficient with sideslip and a loss 
of directional stability due to landing gear deployment 
as observed in Figure 15. The asymmetry in the varia- 
tion of pitching and yawing moment coefficients with 
sideslip is due to the existence of single nose gear door 
as discussed earlier. 

The ground effect aerodynamic data for the base- 
line configuration are shown in Figure 16. It is observed 
that the incremental lift and drag coefficients are posi- 
tive whereas the pitching moment increments are nega- 
tive. This is to be expected because in the presence of 
the ground, the strength of the wing tip vortices dimin- 
ishes leading to a general reduction in downwash along 



the wing span. In a similar fashion, the elevons and body 
fiap were also found to be more effective in presence of 
the ground compared to those in free flight as shown in 
Figures 17 and 18. 



0.3 

0.2 

ACl 0.1 



-0.1 

0.06 
0.04 



., ; ^^^^^— ^;;;l ;' ; 




Figure 16. Incremental lift, drag and pitching moment 

coefficients due to baseline configuration 

in ground effect. 



-0.085 



-0.090 - 



-0.095 




5 10 15 

R. deg 



0.098 
0.096 
0.094 
0.092 
0.090 
0.0B8 
0.086 
0.084 
0.082 



r\ s^--io° i 


1 

a, deg 


% - \^ •16 


\i : ^^^--*^; 


W ; T ; 


^^^ ■ 


V^^;^^^^^^^^^ 







1.0 1.5 

h/b 



Figure 14, Effect of landing gear deployment on 
pitching moment coefficient 



-0.070 
-0.075 
-0.080 
Cm -0.085 
-0.090 
-0.095 
-0.100 



0.015 



y 




\ 


..,.; ;.. N 


^ 


■■: 


y 


^Baseiine 






with landing gea 


^ 


V 

t - 


\ 






; j 




-0.015 



Figure 17. Elevon effectiveness in ground effect 
for5^ = -W. 

0.040 r 



AC, 



m.Sy 



0.025 



0.020 ■ 



0.015 




Figure 15. Combined effect of landing gear and 
sideslip at M^ 0.3, a= 72^. 



Figure 18. Body flap effectiveness in ground effect 

-70°. 



f^^'^f^ 



10 



American Institute of Aeronautics and Astronautics 



The ground effect test data was obtained for some 
combinations of angles of attack, sideslip, eleven, body- 
flap and speedbrake deflections. These data were used 
to perform validation tests for the ground effect aerody- 
namic model. An example of this exercise for a- 8°, 
^3=4°, 5, = -10^ 5^^= -10° and 5^^= 75° (nominal) is 
shown in Figure 1 9. It is observed that the lift, drag and 
pitching moment coefficients predicted by the ground 
effect aerodynamic model are within 3 or 4% of the 
ground effect wind tunnel test data for the combination 
of these parameters. However, the differences in the side 
force, rolling and yawing moment coefficients are much 
higher (Figure 19b). 

The wind tunnel test Reynolds numbers for the X 34 
model (based on mean aerodynamic chord) range up to 
2 X 10'', whereas corresponding full scale flight Rey- 
nolds numbers range up to 40 x 10^. The test Reynolds 
numbers match the flight Reynolds numbers only for a 



a = 8MJ = 4^ 6g = -10^ b^ = -15°, 5sb = 75* fnom) 




On 0.090 



0-085 



0.08 
0.07 - 
C„ 0.06 
0.05 - 
0.04 



iP-Ji^ « O * — — w ■■■■» o » 



1.0 1.5 

h/b 



2.0 



2.5 



(a) 



-0.045 

-O.050 

"^-O-OSS 

-0.060 



a=8^P:=4°,5^ = -10'^ 


5bj = -1 5°, 5,5 = 75* (nom) 


■ -^^r^L 


: o Wind turinei data 
: - Ground effect 
_. :., ,9eroctyn$miQ.rnpdfe(_ 


*OOif) 9 Q jo 


o o; ; 




-0,15 - 



-0.25 




2 3 

Mach numtjer 



Figure 20. Pitching moment coefficient at tunnel and 
flight Reynolds numbers for the baseline X-34 vehicle. 

segment of the hypersonic descent. Elsewhere, the flight 
Reynolds numbers are orders of magnitude higher than 
the wind tunnel test Reynolds numbers. To assess the 
impact of this on the pitch trim which is of critical im- 
portance during unpowered decent, LaRC has conduct- 
ed a limited exercise using various computational fluid 
dynamics (CFD) codes. The results of this exercise are 
shown in Figure 20. The CFD results for the tunnel Rey- 
nolds numbers are shown by open symbols and those 
for the flight Reynolds numbers are shown by filled sym- 
bols. The CFD for the tunnel Reynolds numbers at Mach 
1.05 and 1.25 was run with a turbulent boundary layer 
because the test models in the 16-Foot Transonic Tun- 
nel were tripped. It is observed that the CFD results for 
Mach 2.5, 4.6 and 6.0 agree well with the wind tunnel 
test data. However, the CFD for Mach 1 .05 and 1 .25 
predicts about 10% more nosedown pitching moment 
coefficient compared to the wind tunnel test data. Fur- 
ther, as shown in Figure 20, two CFD codes were run at 
Mach 1 .05 for the flight Reynolds numbers with a tur- 
bulent boundary layer. These limited results indicate that 
the Reynolds number still has some influence and the 
flight vehicle is likely to experience a slightly higher 
nosedown pitching moment than predicted by the wind 
tunnel tests and hence the data in the aero database. This 
increment in nose down pitching moment approximate- 
ly corresponds to about 2° of up elevon deflection. How- 
ever, this aspect was not considered in applying the aero- 
dynamic data in the database to the simulation of vari- 
ous flight trajectories presented in this paper. 

Flight SimMl^tion 



(b) 

Figure 19. Validation test for ground effect 
aerodynamic model. 



Several X-34 Design Reference Mission (DRM) tra- 
jectories have been generated in support of the X-34 
flight test program and are used for envelope expansion 
and flight test range planning purposes. In this paper. 



11 
American Institute of Aeronautics and Astronautics 



four of such DRM trajectories are presented. DRM 1 
refers to a typical low Mach powered flight, DRM 2 re- 
fers to the maximum burn Mach 8 flight, DRM 3 refers 
to a no-engine ignition abort, and DRM 4 represents a 
nominal unpowered approach and landing flight. DRM 
1, DRM 2 and DRM 3 were generated using POST [16] 
and these three trajectories do not include the approach 
and landing phases. The DRM 4 trajectory which in- 
cludes landing phase was generated using STEP [ 1 7] and 
makes use of the ground effect aerodynamic data in the 
aero database. In all these simulations, aerodynamic 
uncertainties including Reynolds number effects were 
not considered. Further, Monte Carlo simulations incor- 
porating aerodynamic and other uncertainties are not 
discussed in this paper. Such simulations are currently 
underway in support of the flight certification program. 

The DRM 1 is representative of the first powered 
(low Mach number) flight of the X-34 vehicle. After 
separation from the L- 1 Oil , the vehicle begins a pull up 



to engine ignition attitude. The engine is ignited and the 
vehicle continues a 2g pull up maneuver. The maximum 
dynamic pressure attained during this flight is about 600 
Ib/sqft. The engine burn is cutoff at a point when about 
50% propellants are still remaining in the tanks. At this 
point, the vehicle dumps the remaining propellants and 
glides back to execute a standard approach and landing. 

The variations of the trajectory parameters for DRM 
1 are presented in Figures 21 to 24. A three dimensional 
plot of the flight trajectory in terms of altitude, down 
range and cross range is given in Figure 21. The maxi- 
mum altitude reached is about 1 15,000 ft, the maximum 
Mach number reached is about 3.6 and the angle of at- 
tack goes up to about 14^^ during the pull up following 
the drop as observed in Figure 22. The time histories of 
the control surface deflections are shown in Figure 23. 
The thrust vectoring (gimbal angle) of about 15° in pitch 
plane is commanded initially during the ascent to aug- 
ment the pitch control. The commanded elevon deflec- 



100 
Down range, nm 




-60 -^ -^^ 

Cross range, nm 



40 60 



tteg -5 



:.\ ; 

I \ I 



Elevon (5^) 
Body flap (6trf) 
Engine ginnbal angle 




50 100 150 200 250 300 350 400 450 500 
Time, sec 



Figure 21. Variation of altitude, down range and cross Figure 23. Time histories of control surface deflection 

range for DRM I. for DRM I. 



Altitude. 



Mach 2 



a, 
deg 




430 



too 200 300 400 500 600 700 800 900 1000 
Time, sec 



eg 

position, 410 - 
in. 




100 200 300 400 

Time, sec 



500 



Figure 22. Time histories of altitude, Mach number Figure 24. Variation of center of gravity during flight 

and angle of attack for DRM L for DRM I. 

12 
American Institute of Aeronautics and Astronautics 



tions reach about -20° when the vehicle is descending 
around Mach 3. With the full scale vehicle likely to ex- 
perience more nose down pitching motnent that approx- 
imately needs an additional -2° elevon deflection to trim 
as discussed earlier, the actual commanded elevon de- 
flection could be about -22°. Although these values of 
elevon deflection are significantly high, they are still 
within the permissible limits. The commanded body flap 
deflections go up to -7.5° during the initial part of the 
ascent and for the rest of the trajectory the body flap 
deflection remains at -10°. The center-of-gravity varia- 
tion is presented in Figure 24. The center-of- gravity 
(e.g.) position at drop is about 404 in from the nose of 
the vehicle. Initially the e.g. moves aft to about 430 in 
and then moves forward to about 393 in and then again 
back to about 417 in. Subsequently, the e.g. remains at 
that position. This pattern of center of gravity movement 
is due to the manner in which LOX is consumed during 
the flight. The LOX is consumed first from the forward 
tank causing the e.g. to move aft. The subsequent for- 
ward shift followed by another rearward movement and 



remaining constant around 417 in is on account of se- 
quential RP and LOX dump. 

The DRM 2 is representative of a full engine burn 
to propellant depletion and vehicle reaching the desig- 
nated altitude of 250,000 ft and target speed of Mach 8. 
The sequence of separation, engine ignition, and pull up 
are similar to the DRM 1 . During this flight, the vehicle 
spends some time outside the atmosphere (dynamic pres- 
sure less than 1 psf) and performs an entry at 25° angle- 
of-attack. The RCS is used during the high altitude flight 
for lateral/directional control. The vehicle then follows 
the standard approach and landing flight path. Stagna- 
tion temperatures during entry can reach 20(X)°F. Enve- 
lope expansion flights will fill the gap between the low 
Mach DRM 1 flight and the maximum Mach 8 DRM 2 
mission. 

The variations of the trajectory parameters for DRM 
2 are presented in Figures 25 to 28. A three dimensional 
plot of the altitude, down and cross ranges is given in 



<105. 



3^ 


' . .. ':' ; .■• ■'' : '"■■. ' ;'■■ . ; I '" ■ . 


2, 


''\ \.- ■■"■ : , . ^ :; ' ' ":-^^^ax a(t J ': " ' 


Attitude, ft 

1, 


I ■ "' " ■• \ VBumout 


0. 
600 


''...,.. ■■■■" Landing r^' ' \ /'■..,: 
500^--^,^^ ■ . . . . .. ■ ^\^ ^Drog,/^200 


Down rj 


mge, nm 2o5''\>^.- ' ^,^^'^'^00 Cross range, nm 
100\„.-^200 
-300 



deg 



Etevon(5g) 

Body flap (5b,) 

Engine gimbal angle 




100 200 300 400 500 600 700 800 900 1000 
Time, sec 



Figure 25, Variation of altitude, down range and cross 
range for DRM 2. 



Figure 27. Time histories of control surface deflection 
for DRM 2. 



. x105 




400 600 1000 

Time, sec 



t400 



e.g. 

position, 

in. 




100 200 300 400 500 600 700 800 900 1000 
Time, sec 



Figure 26. Time histories of altitude, Mach number 
and angle of attack for DRM 2. 



Figure 28. Variation of center of gravity during flight 
for DRM 2. 



13 



American Institute of Aeronautics and Astronautics 



Figure 25. The vehicle attains its target altitude of 
250,000 ft and target speed of Mach 8 around 220 sec- 
onds and then starts its unpowered descent with an an- 
gle of attack of about 25^ as shown in Figure 26. The 
commanded elevon deflections reach about -16° while 
the vehicle is passing through supersonic/transonic 
speeds. As in DRM I, the e.g. moves aft initially due to 
consumption of LOX from forward tank and then for- 
ward due to depletion of aft LOX tank. It then remains 
at about 414 in when all the propellants are depleted and 
engine burn out occurs. 

The DRM 3 is an abort trajectory to deal with en- 
gine failures. Should the main engine fail to ignite after 
separation, a DRM 3 abort mission would be initiated in 
which propellants are immediately dumped and an ap- 
proach and landing to the abort site is conducted. As the 
full propellant load is dumped, the center-of-gravity can 
vary greatly. The DRM 3 abort mission is not a planned 
flight, but would only occur in the case of engine igni- 
tion failure. 

The variations of trajectory parameters for DRM 3 
are shown in Figures 29 to 31. The altitude and Mach 
number steadily decrease following the initiation of the 
abort maneuver as shown in Figure 29. The commanded 
elevon deflections reach up to -20° towards the end as 
shown in Figure 30, As said before for DRM 2, even 
though these elevon deflections are significantly high, 
they are still within permissible limit. The commanded 
speedbrake deflections reach up to 80^ at the beginning 
and towards the end of this mission. Note that the speed- 
brake deflections were not commanded during DRM 1 
and DRM 2. The variation of the e.g. is shown in Figure 
31. The initial aft movement followed by the forward 
movement and then remaining constant around 420 in 
are all caused by the sequential dumping of the RP and 
LOX. 

The unpowered approach and landing test (DRM 4) 
will constitute the first unpowered flight of the X-34 ve- 
hicle. After release from the L- 1 01 1 , the unfueled X-34 
acquires the approach flight path and conducts a stan- 
dard approach and landing. The variation of the trajec- 
tory parameters for DRM 4 are shown in Figures 32 and 
33. It is observed that the vehicle lands around an angle 
of attack of 8*^. The commanded elevon, body flap and 
speedbrake deflections are within limits as in DRM 1 to 
DRM 3. 



clO* 



Altitude, 






Mach 0.5 - 



a, 
deg 




50 100 150 200 250 350 350 400 450 

Time, sec 



Figure 29. Time histories of altitude, Mach number 
and angle of attack for DRM J. 




(Jeg 
-10 ' 




Figure 30. Time histories of control surface deflection 
for DRM 3, 



position, 425 - 
in. 




Figure 31. Variation of center of gravity during flight 
for DRM 3. 



14 
American Institute of Aeronautics and Astronautics 



Altitude. 



Velocity, ^u 
ft/sec ^fi 



Mach 0.5 



a. 
deg 




Figure 32, Time histories of altitude, velocity, Mach 
number and angle of attack for DRM 4. 



CONCLUDING REMARKS 

This paper has presented an overview of the aero- 
dynamic characteristics, the development of the preflight 
aerodynamic database and flight simulations of the 
NASA/Orbital X-34 vehicle. The aerodynamic data in 
the database is provided for both free flight and flight in 
ground effect and covers the complete range of Mach 
numbers, angles of attack, sideslip and control surface 
deflections expected in the entire flight envelope of the 
X-34 vehicle. The variations of the trajectory parame- 
ters and control time histories for four design reference 
missions which arc representative of the X-34 flight test 
program indicate that the vehicle performs these mis- 
sions satisfactorily and the commanded control deflec- 
tions are within the permissible limits at all points along 
these flight trajectories. 

ACKNOWLEDGEMENT 

Authors are thankful to Jim Weilmuensler, Ken 
Sutton, Peter Buning, Ram Prabhu and Shahyar Pirza- 
deh of NASA Langley for the CFD calculations. 

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1 ) Freeman, D.C. Jr., Talay, T.A., and Austin R.E.: 
Reusable Launch Vehicle Technology Program, lAF 96- 
V.4.01, October 1996. 



deg 



5b.. 
deg 





i ^ : rr.::::: 





80 
60 
40 
20 

-20 




; \ 




' ^ • : .. 1 : 


s 


r^ - :~ 




*sb' 


>-' 




Jey 


I_^ ^ 






\ 'i 


i i 



150 



Time, sec 



2) NASA: Reusable Launch Vehicle (RLV), Small 
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Arcachon, France. 



Figure 33, Time histories of control surface deflection 
for DRM 4, 



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15 
American Institute of Aeronautics and Astronautics 



7) Capone, FJ., Bangert, L,S., Asbury, S.C, Mills, 
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