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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 397 records · Page 22
All sourcesNASA NTRSDOE OSTI
NASA NTRS1975

Results of a Pressure Loads Investigation on a 0.030-scale Model (47-OTS) of the Integrated Space Shuttle Vehicle Configuration 5 in the NASA Ames Research Center 11 by 11 Foot Leg of the Unitary Plan Wind Tunnel (IA81A), Volume 1

Results of wind tunnel tests on a 0.030-scale model of the integrated space shuttle vehicle configuration 5 are presented. Testing was conducted in the NASA Ames Research Center 11 x 11 foot leg of the Unitary Plan Wind Tunnel to investigate pressure distributions for airloads analyses at Mach numbers from 0.9 through 1.4. Angles of attack and sideslip were varied from -6 to +6 degrees.

Chee, E.↗
NASA NTRS1976

Results of an investigation to determine local flow characteristics at the air data probe locations using an 0.030-scale model (45-0) of the space shuttle vehicle orbiter configuration 140A/B (modified) in the NASA Ames Research Center unitary plan wind tunnel (OA161, A, B, C), volume 1

Results are presented of wind tunnel test 0A161 of a 0.030-scale model 45-0 of the configuration 140A/B (modified) space shuttle vehicle orbiter in the NASA Ames Research Center Unitary Plan Wind Tunnel facilities. The purpose of this test was to determine local total and static pressure environments for the air data probe locations and relative effectiveness of alternate flight-test probe configurations. Testing was done in the Mach number range from 0.30 to 3.5. Angle of attack was varied from -8 to 25 degrees while sideslip varied between -8 and 8 degrees.

Nichols, M. E.↗
NASA NTRS1992

The Enabler: A reevaluation of design concepts and construction of a scaled model

The basic objective of the student's work this quarter was to make an in depth examination of the design concepts used on the lunar vehicle 'The Enabler'. Several changes were made to the vehicle including a redesigned wheel, a more compact boom and a reduced articulation angle. The vehicle's final dimensions were determined through an optimization process by defining mathematical equations for several of the vehicle's defined objectives. These included the ability to scale a one meter object, traverse a one meter crevice, and maintain a wheel-to-wheel clearance of three inches while at maximum articulation. The final dimensions of the vehicle were used to construct an approximate 1/4 scale model of the chassis and wheels. The boom, however, was constructed on a 1/5 scale (from the original design). This was due to the redesign of the boom and the limitations of the constructing material and PVC fittings.

Bess, David↗
NASA NTRS1973

Aerodynamic characteristics of a large-scale model with a swept wing and a jet flap having an expandable duct

The data from an investigation of the aerodynamic characteristics of the expandable duct-jet flap concept are presented. The investigation was made using a large-scale model in the Ames 40- by 80-foot Wind Tunnel. The expandable duct-jet flap concept uses a lower surface, split flap and an upper surface, Fowler flap to form an internal, variable area cavity for the blowing air. Small amounts of blowing are used on the knee of the upper surface flap and the knee of a short-chord, trailing edge control flap. The bulk of the blowing is at the trailing edge. The flap could extend the full span of the model wing or over the inboard part only, with blown ailerons outboard. Primary configurations tested were two flap angles, typical of takeoff and landing; symmetric control flap deflections, primarily for improved landing performance; and asymmetric aileron and control flap deflections, for lateral control.

Aiken, T. N.↗
NASA NTRS1974

Results of tests in the NASA/LaRC 31 inch CFHT on an 0.010-scale model (32-OT) of the space shuttle configuration 3 to obtain hypersonic aerodynamic characteristics for second stage operation during nominal boost and the abort RTLS mode (IA58)

Tests were conducted to obtain hypersonic aerodynamic forces and moments on an 0.010-scale model of the space shuttle vehicle configuration 3. Hypersonic stability data were obtained from tests at Mach 10.3 and dynamic pressure of 150 psf for the integrated orbiter and external tank, orbiter alone, and external tank alone. The effects of solid plume simulation from the main propulsion system as well as elevon, aileron, and rudder deflections were also investigated.

Thornton, D. E.↗
NASA NTRS1975

Aerodynamic results of a separation effects test on a 0.01-scale model (52-OTS) of integrated SSV in the AEDC/VKF 40-by-40 inch supersonic wind tunnel A, volume 1

Experimental aerodynamic investigations were conducted, during the period July 18-19, 1974, in the AEDC/VKF Tunnel A facility on a 0.01-scale model (52-OTS) of the integrated space shuttle vehicle, including only one SRB. The purpose of the investigation was to obtain data for close-in proximity (SRB to orbiter/tank) effects with the orbiter/tank combination at relatively high alpha and beta attitudes, and with the SRB separation motors off. The AEDC Captive Trajectory System (CTS), which supported the SRB, was used in conjunction with the tunnel primary sector (supporting the orbiter/tank) to obtain grid type separation effects data. The one symmetrical SRB model was used interchangeably to obtain both right-hand and left-hand SRB data. Free-stream data were also obtained for the orbiter/tank and for the SRB. This data was used to provide baselines for proximity effects. The entire investigation was conducted at a free-stream Mach number of 4.5 with unit Reynolds number ranging from 4.0 to 6.5 million per foot.

Campbell, J. H., II↗
NASA NTRS1976

A low speed wind tunnel test of a 0.050 scale model of shuttle orbiter (model 089B) to investigate the longitudinal and lateral directional effects of canard and tail configurational modifications in the LTV LSWT (MA14)

A low speed wind tunnel test was conducted to determine the effects of 6 canard configurations on the 0.050 scale model of shuttle orbiter 089B. In addition, two horizontal tail configurations were tested at two positions on the model as were two wing configurations. Since this test was restricted to 103 runs, only a limited number of permutations of the configurational changes could be tested. The testing was done in the 15 by 20 foot section of the LSWT and consisted of pitch polars, one yawed polar and several yaw runs. The pitch polars encompassed an alpha range from 0 to 28 deg; the yawed polar was run at beta = +2 degrees and the yaw runs covered a beta range from -6 to +6 deg at angles-of-attack of 0, 4, 10, 16, and 20 deg.

Chambliss, E. B.↗
NASA NTRS1947

Flight-Test Evaluation of the Longitudinal Stability and Control Characteristics of 0.5-Scale Models of the Fairchild Lark Pilotless-Aircraft Configuration: Standard Configuration with Wing Flaps Deflected 60 Degrees and Model having Tail in Line with Wings, TED No. NACA 2387

Flight tests were conducted at the Flight Test Station of the Pilotless Aircraft Research Division at Wallop Island, Va., to determine the longitudinal control and stability characteristics of 0.5-scale models of the Fairchild Lark pilotless aircraft with the tail in line with the wings a d with the horizontal wing flaps deflected 60 deg. The data were obtained by the use of a telemeter and by radar tracking.

Stone, David G.↗
NASA NTRS1956

Longitudinal and Lateral Stability, Control Characteristics, and Vertical-Tail-Load Measurements for 0.03-Scale Model of the Avro CF-105 Airplane at Mach Number 1.41

An investigation has been made in the Langley 4- by 4-foot supersonic pressure tunnel at a Mach number of 1.41 to determine the aerodynamic characteristics of an 0.03-scale model of the Avro CF-105 airplane. The investigation included the determination of the static longitudinal and lateral stability, the control and the hinge-moment characteristics of the elevator, the aileron, and the rudder, as well as the vertical-tail-load characteristics. The results indicated a minimum drag coefficient of about 0.0270, and a maximum trimmed lift-drag ratio of about 4.25 which occurs at a lift coefficient of 0.16. The directional stability decreased with increasing angle of attack until a region of static instability occurred above an angle of attack of about 9 deg.

Spearman, M. Leroy↗
NASA NTRS2005

Acoustic and Aero-Mixing Experimental Results for Fluid Shield Scale Model Nozzles

The principle objectives of this investigation are to evaluate the acoustic and aerodynamic characteristics of fluid shield nozzle concept and to assess Far 36, Stage 3 potential for fluid shield nozzle with Flade Cycle. Acoustic data for nine scale model nozzle configurations are obtained. The effects of simulated flight and geometric and aerothermodynamic flow variables on the acoustic behavior of the fluid shield are determined. The acoustic tests are aimed at studying the effect of: (1) shield thickness, (2) wrap angle, (3) mass flow and velocity ratios between shield and core streams at constant cycle specific thrust (i.e., mixed velocity), (4) porous plug, and (5) subsonic shield. Shadowgraphs of six nozzle configurations are obtained to understand the plume flowfield features. Static pressure data on suppressor chutes in the core stream (shielded and unshielded) sides and on plug surface are acquired to determine the impact of fluid shield on base drag of the 36-chute suppressor nozzle and the thrust augmentation due to the plug, respectively.

Salikuddin, M.↗
NASA NTRS1975

Results of a jet plume effects test on Rockwell International integrated space shuttle vehicle using a vehicle 5 configuration 0.02-scale model (88-OTS) in the 11 by 11 foot leg of the NASA/Ames Research Center unitary plan wind tunnel (IA19), volume 1

Results are presented of jet plume effects test IA19 using a vehicle 5 configuration integrated space shuttle vehicle 0.02-scale model in the NASA/Ames Research Center 11 x 11-foot leg of the unitary plan wind tunnel. The jet plume power effects on the integrated vehicle static pressure distribution were determined along with elevon, main propulsion system nozzle, and solid rocket booster nozzle effectiveness and elevon hinge moments.

Nichols, M. E.↗
NASA NTRS1948

High-Speed Wind-Tunnel Investigation of the Longitudinal Stability and Control Characteristics of a 0.10-Scale Model of the Grumman XF9F-2 Airplane, TED No. NACA DE301

An investigation was made in the Langley high-speed 7-by 10-foot tunnel to determine the high-speed longitudinal stability end con&o1 characteristics of a 0.01-scale model of the Grumman XF9F-2 airplane in the Mach number range from 0.40 to 0.85. The results indicated that the lift and drag force breaks occurred at a Mach number of about 0.76. The aerodynamic-center position moved rearward after the force break and control position stability was present for all Mach numbers up to a Mach number of 0.80.

Polhamus, Edward C.↗
NASA NTRS1949

High-Speed Wind-Tunnel Investigation of the Lateral Stability Characteristics of a 0.10-Scale Model of the Grumman XF9F-2 Airplane, TED No. NACA DE 301

An investigation was made in the Langley high-speed 7- by 10-foot tunnel to determine the high-speed lateral and directional stability characteristics of a 0.10-scale model of the Grumman XF9F-2 airplane in the Mach number range from 0.40 to 0.85. The results indicate that static lateral and directional stability is present throughout the Mach number range investigated although in the Mach number range from 0.75 to 0.85 there is an appreciable decrease in rolling moment due to sideslip. Calculations of the dynamic stability indicate that according to current flying-quality requirements the damping of the lateral oscillation, although probably satisfactory for the sea-level condition, may not be satisfactory for the majority of the altitude conditions investigated

Polhamus, Edward C.↗
NASA NTRSDate not supplied

Learjet Update Flight and Scale-Model-Nozzle Acoustics Test

The impetus for the work is the need for an improved ability to predict takeoff noise of future supersonic commercial transport aircraft. The overall objective is to validate as well as improve the conversion of facility scale-model data and semi-empirical jet-noise predictions to flight noise. Acoustic data from a Learjet-25 flyover test and a complementary rig-test are compared. The uncertainty in the rig-to-flight transformation is presented. Ultimately, the goal is improved noise-prediction methods for system studies of future commercial supersonic aircraft.

Lennart S Hultgren↗
NASA NTRS1987

Experimental Evaluation of Turning Vane Designs for High-speed and Coupled Fan-drive Corners of 0.1-scale Model of NASA Lewis Research Center's Proposed Altitude Wind Tunnel

Two turning vane designs were experimentally evaluated for the fan-drive corner (corner 2) coupled to an upstream diffuser and the high-speed corner (corner 1) of the 0.1 scale model of NASA Lewis Research Center's proposed Altitude Wind Tunnel. For corner 2 both a controlled-diffusion vane design (vane A4) and a circular-arc vane design (vane B) were studied. The corner 2 total pressure loss coefficient was about 0.12 with either vane design. This was about 25 percent less loss than when corner 2 was tested alone. Although the vane A4 design has the advantage of 20 percent fewer vanes than the vane B design, its vane shape is more complex. The effects of simulated inlet flow distortion on the overall losses for corner 1 or 2 were small.

Gelder, Thomas F.↗
NASA NTRS1946

Flight-Test Evaluation of the Longitudinal Stability and Control Characteristics of 0.5-Scale Models of the Fairchild Lark Pilotless-Aircraft Configuration. Model with Wing Flaps Deflected 15 Degrees, TED No. NACA 2387

A flight test was conducted at the Flight Test Station of the Pilotless Aircraft Research Division at Wallops Island, Va., to determine the longitudinal control and stability characteristics of a 0.5-scale model of the Fairchild Lerk Pilotless aircraft with the horizontal wing flaps deflected 15 degrees. The data were obtained by the use of a telemeter and also by radar tracking. The results show an increase of effectiveness of the longitudinal control in producing normal accelerations up to a Mach number of 0.75 where this effectiveness gradually decreased becoming negative at a Mach number of 0.89. Previous tests with wing flaps undeflected an increase in effectiveness up to Mach number of 0.93 where a sudden loss of control occurred. The model was dynamically stable throughout the speed range. The data confirmed the drag increase at the critical Mach number for large angles of attack is indicated in high-speed wind-tunnel tests.

Stone, David G.↗
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.↗