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At least 1,711 records · Page 95
All sourcesNASA NTRSDOE OSTI
NASA NTRS1989

Calculation of real-gas effects on blunt-body trim angles

The effect of vibrational excitation and dissociation at high temperatures on the trim angle of attack of a blunt lifting body is calculated for a nonequilibrium flow regime in air using a CFD technique. The vibrational-electronic temperature and the species densities are calculated assuming the flow to be in a nonequilibrium state. The forebody flow of a two-dimensional blunt body of the shape of the Apollo Command Module at a finite angle of attack is calculated. The results show that the pitching moment around a reference point is larger and the trim angle of attack is smaller for a reacting gas than for a perfect gas. The calculated shift in the trim angle due to the real-gas effect is of the same order as that seen during the Apollo flights.

Park, Chul↗
NASA NTRS1972

Wind-tunnel studies of wing wake turbulence.

Description of velocity measurements made in the wake of wings in the Ames 7 x 10 ft wind tunnel. Distributions of velocity components were measured with a three-wire anemometer up to 12 chord lengths downstream of a CV-990 aircraft model and a rectangular wing. Results show that increasing the drag increases the vortex core radius, reduces the maximum tangential velocities, and increases the magnitude of axial velocity defects. For the rectangular wing, axial velocity changes from a defect (wake flow) for angles of attack less than 9 deg to an excess (jet flow) for angles of attack greater than 9 deg. Wind-tunnel measurements of the near flowfield are compared with flight measurements of the far flowfield.

Chigier, N. A.↗
NASA NTRS1972

Space shuttle: Static surface pressures of the 0.004 scale 049 orbiter in the launch configuration

Wing and lower body surface static pressure data for the space shuttle 049 orbiter while in the launch configuration were obtained. The purpose of the test was to determine the optimum incidence position of the orbiter relative to the hydrogen-oxygen (HO) tank and the optimum radial position of the solid rocket motors (SRM) on the HO tank. The orbiter was mounted on the HO tank at incidence angles of 0 and -1.5 degrees to determine the optimum incidence position. The SRM boosters were tested at radial positions of 75, 90, and 135 degrees on the HO tank to determine their optimum position with respect to the loads imposed on the orbiter. The test was conducted in the tunnel over a Mach number range of 0.6 to 4.96. Angle of attack was varied from -8 to +8 degrees at zero degree angle of sideslip, and at sideslip angles varying from -6 to +6 degrees at zero degree angle of attack.

Buchholz, R. E.↗
NASA NTRS1973

Results of investigations on a 0.015-scale model 2A configuration of the Rockwell International space shuttle orbiter in the NASA/Ames Research Center 3.5 foot hypersonic wind tunnel

Experimental aerodynamic investigations were conducted in the NASA/Ames 3.5-Foot Hypersonic wind Tunnel during the interim April 9-18, 1973 on a 0.015-scale model of the Rockwell International Space Shuttle Orbiter, configuration 2A. Six component aerodynamic force and moment data were recorded over an angle of attack range from -3 deg to 42 deg at 0 deg angle of sideslip and from -10 deg to 10 deg sideslip at 0 deg and 45 deg constant angle of attack. Test Mach numbers were 5.27 and 7.32 at unit Reynolds number of 2.5 million per foot. Various elevon, rudder, speedbrake, and body flap deflections were tested to determine longitudinal and lateral-directional stability characteristics and to establish trim capability.

Milam, M. D.↗
NASA NTRS1974

Results of investigations on an 0.015 scale 140A/B configuration space shuttle vehicle orbiter model (49-0) in the LTV 4 by 4-foot high speed wind tunnel (0A84)

Data obtained from a wind tunnel test of an 0.015-scale 140 A/B configuration space shuttle vehicle orbiter model (49-0) are presented. Runs were conducted at Mach numbers 0.6, 0.9, 1.2, 1.6, and 2.0 and Reynolds numbers of 8.4, 9.0, 9.0, 10.9, and 13.6 million per foot respectively. Various control surface settings were investigated from angles of attack of minus 4 degrees to plus 30 degrees at fixed angles of sideslip of zero and minus 5 degrees and through angles of sideslip from minus 2 to plus 8 degrees at fixed angles of attack of zero, plus 10, plus 15, and plus 20 degrees. The purpose of the test was to define the longitudinal and lateral-directional stability and control characteristics for the updated SSV configuration.

Esparza, V.↗
NASA NTRS1975

Results of investigations on a 0.004-scale 140C modified configuration space shuttle vehicle orbiter model (74-0) in the NASA/Langley Research Center hypersonic helium tunnel

Data obtained during a wind tunnel test of a 0.004-scale 140C modified configuration SSV orbiter are documented. The test was conducted during August 1974 with 80 occupancy hours charged, and all runs were conducted at a nominal Mach number of 20 and at Reynolds numbers of 0.7, 1.0, 1.8, and 1,100,000 based on body length. The complete -140C modified model was tested with various elevon settings at angles of attack from 10 to 50 degrees at zero yaw and from angles of sideslip of -10 to +10 at 35 deg angle of attack. The purpose of this test was to obtain high hypersonic longitudinal and lateral-directional stability and control characteristics of the updated SSV configuration.

Hawthorne, P. J.↗
NASA NTRS1977

Aerodynamic characteristics at Mach numbers from 0.33 to 1.20 of a wing-body design concept for a hypersonic research airplane

An experimental investigation of the static aerodynamic characteristics of a model of one design concept for the proposed National Hypersonic Flight Research Facility was conducted in the Langley 8 foot transonic pressure tunnel. The experiment consisted of configuration buildup from the basic body by adding a wing, center vertical tail, and a three module or six module scramjet engine. The freestream test Mach numbers were 0.33, 0.80, 0.90, 0.95, 0.98, 1.10, and 1.20 at Reynolds numbers per meter ranging from 4.8 x 1 million to 10.4 x 1 million. The test angle of attack range was approximately -4 deg to 22 deg at constant angles of sideslip of 0 deg and 4 deg; the angle of sideslip ranged from about -6 deg to 6 deg at constant angles of attack of 0 deg and 17 deg. The elevons were deflected 0 deg, -10 deg, and -20 deg with rudder deflections of 0 deg and 15.6 deg.

Dillon, J. L.↗
NASA NTRS1979

Some theoretical considerations of a stall proof airplane

For the stall proof airplane there should be a stabilizing pitching moment below the stall angle of attack of sufficient amount to prevent the attainment of the stall angle of attack which cannot be over-ridden by control deflections. This paper presents (1) a development of the moment equations to show the theoretical considerations of a stall proof airplane and (2) the nonlinear moment characteristics that must be obtained to satisfy the stall proof requirements. In addition, it is shown that an aerodynamic spoiler located on the under surface of the horizontal tail can be designed to meet these requirements of a stall proof airplane. Wind tunnel results are shown to validate assumptions and predictions.

Chevalier, H. L.↗
NASA NTRS1979

Aerodynamic interactions from reaction controls for lateral control of the M2-F2 lifting-body entry configuration at transonic and supersonic and supersonic Mach numbers

Tests were conducted in the Ames 6 by 6 foot wind tunnel to determine the interaction of reaction jets for roll control on the M2-F2 lifting-body entry vehicle. Moment interactions are presented for a Mach number range of 0.6 to 1.7, a Reynolds number range of 1.2 x 10 to the 6th power to 1.6 x 10 to the 6th power (based on model reference length), an angle-of-attack range of -9 deg to 20 deg, and an angle-of-sideslip range of -6 deg to 6 deg at an angle of attack of 6 deg. The reaction jets produce roll control with small adverse yawing moment, which can be offset by horizontal thrust component of canted jets.

Bailey, R. O.↗
NASA NTRS1980

Effect of wing location and strakes on stability and control characteristics of a monoplanar circular missile with low-profile tail fins at supersonic speeds

A wind tunnel test was conducted at Mach numbers from 1.70 to 2.86 to extend the aerodynamic data base for wing tail effect on stability and control characteristics of monoplanar missiles. The results are summarized to show the effects of tail fin dihedral angle, wing location, and nose body strakes. The results indicate that an increase in tail fin dihedral angle produces positive increments in directional stability that allow greater trimmed lift coefficient values (maneuver potential) to be obtained. An increase in wing tail gap for the Mach number range reduces the aerodynamic center travel and produces reductions in directional stability at the lower angles of attack. A change in wing height (vertical location) strongly influences the angle of attack at which pitch up and the most directional stability occur. The addition of strakes to the baseline configuration increases directional stability, which allows a significant increase in stable trimmed maneuver capability. The tail fins of the baseline configuration are effective in producing roll and yaw control that are accompanied by favorable yaw and roll, respectively.

Blair, A. B., Jr.↗
NASA NTRS1983

Augmentation of fighter-aircraft performance by spanwise blowing over the wing leading edge

Spanwise blowing over the wing and canard of a 1:35 model of a close-coupled-canard fighter-airplane configuration (similar to the Kfir-C2) was investigated experimentally in low-speed flow. Tests were conducted at airspeeds of 30 m/sec (Reynolds number of 1.8 x 10 to the 5th power based on mean aerodynamic chord) with angle-of-attack sweeps from -8 deg to 60 deg, and yaw-angle sweeps from -8 deg to 36 deg at fixed angles of attack 0 deg, 10 deg, 20 deg, 25 deg, 30 deg, and 35 deg. Significant improvement in lift-curve slope, maximum lift, drag polar and lateral/directional stability was found, enlarging the flight envelope beyond its previous low-speed/maximum-lift limit. In spite of the highly swept (60 deg) leading edge, the efficiency of the lift augmentation by blowing was relatively high and was found to increase with increasing blowing momentum on the close-coupled-canard configuration. Interesting possibilities of obtaining much higher efficiencies with swirling jets were indicated.

Seginer, A.↗
NASA NTRS1984

Wing design for spin resistance

Use of a discontinuous outboard wing leading edge to improve stall/spin characteristics has been evaluated through wind-tunnel and flight tests. Addition of such a discontinuous outboard wing leading-edge droop design to three light airplanes having NACA 6-series airfoil sections produced significant improvements in stall characteristics and spin resistance. The increased spin resistance of the modified airplanes has been related to the difference in angle of attack between the outer wing panel stall and the maximum attainable angle of attack.

Stough, H. P., III↗
NASA NTRS1985

Measurements of canard-induced roll oscillations

A small canard wing was installed in front of a delta wing which was mounted on a free-to-roll sting balance in a low-speed wind tunnel. The leading edge vortices, originating from the canard, enhanced self-induced roll oscillations at test conditions for which the basic delta wing would otherwise have been stable. Time-dependent roll angle, and normal and side force data recorded during these oscillations are presented along with their phase relations. It was found that the canard increased the range of angle of attack at which self-induced oscillations occurred. Also, at an angle of attack of about 46 deg asymmetric oscillations are observed.

Katz, J.↗
NASA NTRS1985

Leading-edge design for improved spin resistance of wings incorporating conventional and advanced airfoils

Discontinuous wing leading-edge droop designs have been evaluated as a means of modifying wing autorotative characteristics and thus improving airplane spin resistance. Addition of a discontinuous outboard wing leading-edge droop to three typical light airplanes having NACA 6-series wing sections produced significant improvements in stall characteristics and spin resistance. Wind tunnel tests of two wings having advanced natural laminar flow airfoil sections indicated that a discontinuous leading-edge droop can delay the onset of autorotation at high angles of attack without adversely affecting the development of laminar flow at cruise angles of attack.

Stough, H. P., III↗
NASA NTRS1988

Application of Navier-Stokes analysis to stall flutter

A solution procedure was developed to investigate the two-dimensional, one- or two-dimensional flutter characteristics of arbitrary airfoils. This procedure requires a simultaneous integration in time of the solid and fluid equations of motion. The fluid equations of motion are the unsteady compressible Navier-Stokes equations, solved in a body-fitted moving coordinate system using an approximate factorization scheme. The solid equations of motion are integrated in time using an Euler implicit scheme. Flutter is said to occur if small disturbances imposed on the airfoil attitude lead to divergent oscillatory motions at subsequent times. The flutter characteristics of airfoils in subsonic speed at high angles of attack and airfoils in high subsonic and transonic speeds at low angles of attack are investigated. The stall flutter characteristics are also predicted using the same procedure.

Wu, J. C.↗
NASA NTRS1989

Surface pressure distributions on a delta wing undergoing large amplitude pitching oscillations

Wind tunnel experiments were performed on a 70 deg sweep delta wing to determine the effect of a sinusoidal pitching motion on the pressure field on the suction side of the wing. Twelve pressure taps were placed from 35 to 90 percent of the chord, at 60 percent of the local semi-span. Pressure coefficients were measured as a function of Reynolds number and pitch rate. The pressure coefficient was seen to vary at approximately the same frequency as the pitching frequency. The relative pressure variation at each chord location was comparable for each case. The average pressure distribution through each periodic motion was near the static distribution for the average angle of attack. Upon comparing the upstroke and downstroke pressures for a specific angle of attack, the downstroke pressures were slightly larger. Vortex breakdown was seen to have the most significant effect at the 40 to 45 percent chord location, where a decrease in pressure was apparent.

Thompson, Scott A.↗
NASA NTRS1989

Longitudinal stability and control characteristics of the Quiet Short-Haul Research Aircraft (QSRA)

Flight experiments were conducted to evaluate various aerodynamic characteristics of the Quiet Short-Haul Research Aircraft (QSRA), an experimental aircraft that makes use of the upper-surface blown (USB) powered-lift concept. Time-history records from maneuvers performed with the aircraft in landing-approach and take-off configurations (with its stability augmentation system disengaged) were analyzed to obtain longitudinal stability and control derivatives and performance characteristics. The experiments included measuring the aircraft responses to variations in the deflection of direct-lift control spoilers and to thrust variations as well as to elevator inputs. The majority of the results are given for the aircraft in a landing configuration with the USB flaps at 50 degrees. For this configuration, if the static longitudinal stability is defined as the variation of the pitching-moment coefficient with the lift coefficient at a constant thrust coefficient, this stability decreases significantly with increasing angle of attack above 9 degrees. For this configuration, at small and negative angles of attack and high levels of thrust, the elevators and the horizontal stabilizer lost effectiveness owing to incipent stalling, but this occurred only during unsteady maneuvers and for brief time intervals.

Stephenson, Jack D.↗
NASA NTRS1955

Flight Measurements of Horizontal-Tail Loads on the Bell X-5 Research Airplane at a Sweep Angle of 58.7 Deg

A flight investigation was made at altitudes of 40,000, 25,000 and 15,000 feet to determine the horizontal-tail loads of the Bell X-5 research airplane at a sweep angle of 58.7 deg over the lift range of the airplane for Mach numbers from 0.61 to 1.00. The horizontal-tail loads were found to be nonlinear with lift throughout the lift ranges tested at all Mach numbers except at a Mach number of 1.00. The balancing tail loads reflected the changes which occur in the wing characteristics with increasing angle of attack. The nonlinearities were, in general, more pronounced at the higher angles of attack near the pitch-up where the balancing tail loads indicate that the wing-fuselage combination becomes unstable. No apparent effects of altitude on the balancing tail loads were evident over the comparable lift ranges of these tests at altitudes from 40,000 feet to 15,000 feet. Comparisons of balancing tail loads obtained from flight and windtunnel tests indicated discrepancies in absolute magnitudes, but the general trends of the data agree. Some differences in absolute magnitude may be accounted for by the tail load carried inboard of the strain-gage station and the load induced on the fuselage by the presence of the tail. These loads were not measured in flight.

Reed, Robert D↗