Scientific and Technical Aerospace Reports Volume 38 July 28, 2000

English; 34th; 34th Thermophysics Conference, 19-22 Jun. 2000, Denver, CO, USA; Sponsored by American Inst. of Aeronautics
and Astronautics, USA; Original contains color illustrations
Report No.(s): AIAA Paper 2000-2685; Copyright Waived; Avail: CASI; A03, Hardcopy; A01, Microfiche
The X-38 program seeks to demonstrate an autonomously returned orbital test flight vehicle to support the development of
an operational Crew Return Vehicle for the International Space Station. The test flight, anticipated in 2002, is intended to demonstrate
the entire mission profile of returning Space Station crew members safely back to earth in the event of medical or mechanical
emergency. Integral to the formulation of the X-38 flight data book and the design of the thermal protection system, the aerothermodynamic
environment is being defined through a synergistic combination of ground based testing and computational fluid
dynamics. This report provides an overview of the hypersonic aerothermodynamic wind tunnel program conducted at the NASA
Langley Research Center in support of the X-38 development. Global and discrete surface heat transfer force and moment, surface
streamline patterns, and shock shapes were measured on scaled models of the proposed X-38 configuration in different test gases
at Mach 6, 10 and 20. The test parametrics include angle of attack from 0 to 50 degs, unit Reynolds numbers from 0.3 x 10 (exp
6) to 16 x 10 (exp 6)/ ft, rudder deflections of 0, 2, and 5 deg. and body flap deflections from 0 to 30 deg. Results from hypersonic
aerodynamic screening studies that were conducted as the configuration evolved to the present shape at, presented. Heavy gas
simulation tests have indicated that the primary real gas effects on X-38 aerodynamics at trim conditions are expected to favorably
influence flap effectiveness. Comparisons of the experimental heating and force and moment data to prediction and the current
aerodynamic data book are highlighted. The effects of discrete roughness elements on boundary layer transition were investigated
at Mach 6 and the development of a transition correlation for the X-38 vehicle is described. Extrapolation of ground based heating
measurements to flight radiation equilibrium wall temperatures at Mach 6 and 10 were made and generally compared to within
50 deg F of flight prediction.
Author
Aerothermodynamics; International Space Station; X-38 Crew Return Vehicle; Flight Tests; Computational Fluid Dynamics;
Hypersonic Wind Tunnels; Wind Tunnel Tests
20000064077 NASA Ames Research Center, Moffett Field, CA USA
Comment on ’Guidance for an Aerocapture Maneuver
Tauber, Michael E., NASA Ames Research Center, USA; Journal of Guidance, Control, and Dynamics; July - August 1994; Volume
17, No. 4, pp. 878-878; In English; Sponsored by American Inst. of Aeronautics and Astronautics, USA; Copyright; Avail:
Issuing Activity; Abstract Only
It is stated that the aerodynamic forces on the vehicle being aerocaptured are controlled by ”altering the angle of attack” and
thereby controlling the lift coefficient. Furthermore, the resulting variation of drag coefficient with angle of attack was ignored.
The purpose of this Comment is to point out that an aerodynamic control method that is much more effective than the pitch modulation
has been studied and utilized during entries for many years. During aerocapture, it is desirable to have a large range of lift
coefficients available, while keeping the vehicle’s ballistic coefficients constant. This is accomplished by modulating the vehicle’s
bank angle, i.e., by rolling the vehicle about its velocity vector. by this method, the angle of attack can be held constant (at the
trim angle, if desired), and the C(sub D) and the ballistic coefficient remain constant. Furthermore, the vertical component of the
normal force vector (essentially the lift) can be varied over its entire range, from maximum positive to maximum negative values.
Reaction controls, rather than aerodynamic ones, are usually utilized to change the bank angle of the vehicle, thus requiring the
use of fuel. However, the fuel expenditure that is required to change the bank angle is far less than the amount that would have
to be used to continuously hold the vehicle at pitch angles that differ significantly from its trim angle of attack. Also, it has been
shown that bank angle modulation to vary the lift can enlarge the entry corridor by increasing the entry angle for the undershoot
boundary, where both the heating rate and deceleration reach a maximum. Finally, the crew’s deceleration tolerance can be
increased somewhat when the bank angle is varied, as opposed to the pitch angle. For bank modulation, the deceleration force
vector can be kept at a constant angle with respect to the occupants whose tolerance to g loads is highest when the force is applied
in a direction normal to the upper torso. The advantages of bank angle variation to modulate the lift vector were recognized long
ago, and this method of control was used successfully on the Apollo command module during lunar return’ and, more recently,
for the Space Shuttle Orbiter.
Author
Aerocapture; Aerodynamic Forces; Angle of Attack; Control Surfaces; Pitch (Inclination)
20000064088 Technische Hogeschool Twente, Faculty of Applied Mathematics, Enschede, Netherlands
Performance of an Implicit Algorithm for Inviscid Flow Around an Airfoil
Strating, P.; vanBuuren, R.; Jun. 25, 1997; 22p; In English
Report No.(s): PB2000-104904; No Copyright; Avail: National Technical Information Service (NTIS)
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