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GaAs p(+ +)n(+)n(-)n(+) single-drift tunnel injection transit-time (TUNNETT) diodes for W-band operation have been successfully designed and tested. An output power of 32 mW at 93.5 GHz with a dc to RF conversion efficiency of 2.6 percent was obtained. The oscillations have a clean spectrum in a conventional waveguide cavity.
The use of higher harmonic control (HHC) of blade pitch to reduce blade-vortex interaction (BVI) noise is examined by means of a rotor acoustic test. A dynamically scaled, four-bladed, articulated rotor model was tested in a heavy gas (Freon-12) medium. Acoustic and vibration measurements were made for a large range of matched flight conditions where prescribed (open loop) HHC pitch schedules were superimposed on the normal (baseline) collective and cyclic trim pitch. A novel sound power measurement technique was developed to take advantage of the reverberance in the hard walled tunnel. Quantitative sound power results are presented for a 4/rev (4P) collective pitch HHC. By comparing the results using 4P HHC to corresponding baseline (no HHC) conditions, significant midfrequency noise reductions of 5-6 dB are found for low-speed descent conditions where BVI is most intense. For other flight conditions, noise is found to increase with the use of HHC. LF loading noise, as well as fixed and rotating frame vibration levels, show increased levels.
NASA Langley Research Center, in partnership with NASA Marshall Space Flight Center and NASA Ames Research Center, was involved in the aerodynamic analyses, testing, and database development for the Ares I A106 crew launch vehicle in support of the Ares Design and Analysis Cycle. This paper discusses the development of lift-off/transition and ascent databases. The lift-off/transition database was developed using data from tests on a 1.75% scale model of the A106 configuration in the NASA Langley 14x22 Subsonic Wind Tunnel. The power-off ascent database was developed using test data on a 1% A106 scale model from two different facilities, the Boeing Polysonic Wind Tunnel and the NASA Langley Unitary Plan Wind Tunnel. The ascent database was adjusted for differences in wind tunnel and flight Reynolds numbers using USM3D CFD code. The aerodynamic jet interaction effects due to first stage roll control system were modeled using USM3D and OVERFLOW CFD codes.
Ground static and wind tunnel test of a scale model modified T-39 airplane are reported. The configuration in the nose and replacement of the existing nacelles with tilting lift/cruise fans. The model was powered with three 14 cm diameter tip driven turbopowered simulators. Forces and moments were measured by an internal strain guage balance. Engine simulator thrust and mass flow were measured by calibrated pressure and temperature instrumentation mounted downstream of the fans. The low speed handling qualities and general aerodynamic characteristics of the modified T-39 were defined. Test variables include thrust level and thrust balance, forward speed, model pitch and sideslip angle at forward speeds, model pitch, roll, and ground height during static tests, lift/cruise fan tilt angle, flap and aileron deflection angle, and horizonal stabilizer angle. The effects of removing the landing gear, the lift/cruise fans, and the tail surfaces were also investigated.
Previous aerodynamic heating tests have been conducted in the AEDC/VKF Supersonic Wind Tunnel (A) to aid in defining the design thermal environment for the space shuttle external tank. The quality of these data has been under discussion because of the effects of low tunnel enthalpy and slow model injection rates. Recently the AEDC/VKF Hypersonic Wind Tunnel (C) has been modified to provide a Mach 4 capability that has significantly higher tunnel enthalpy with more rapid model injection rates. Tests were conducted in Tunnel C at Mach 4 to obtain data on the external tank for comparison with Tunnel A results. Data were obtained on a 0.0175 scale model of the Space Shuttle Integrated Vehicle at Re/ft = 4 x 10 to the 6th power with the tunnel stagnation temperature varying from 740 to 1440 R. Model attitude varied from an angle of attack of -5 to 5 deg and an angle of sideslip of -3 to 3 deg. One set of data was obtained in Tunnel C at Re/ft = 6.9 x 10 to the 6th for comparison with flight data. Data comparisons between the two tunnels for numerous regions on the external tank are given.
A JT8D-17 turbofan engine was tested in the NASA-Ames 40- by 80-foot wind tunnel to determine flight effects on jet and fan noise. Baseline, quiet nacelle with 20-lobe ejector/suppressor, and internal mixer configurations were tested over a range of engine power settings and tunnel velocities. Flight effects derived from the 40- by 80-foot wind tunnel test are compared with 727/JT8D flight test data and with model data obtained in a smaller wind tunnel. Procedures are defined for measuring noise data in a wind tunnel relatively near the sources and analyzing the results to obtain far-field flight effects. Wind tunnel and 727 flight test noise results compare favorably for both the baseline and quiet nacelle configurations. Two reports are provided, including a comprehensive version with extensive test results and analysis and the subject summary version that emphasizes data analysis and program finding.
Design features and certain performance data are outlined for three different basic airframes. The heart of the design is a powered lift system known formally as the Advanced Blown Lift Enhancement system. A summary of the results of both powered-model wind tunnel tests of the airplane and of static tests of the vectoring-nozzle system is given to indicate the rather high degree of thrust vectoring efficiency and powered lift enhancement attained.
Theoretical considerations indicate that cooling the wind-tunnel test gas to cryogenic temperatures will provide a large increase in test Reynolds number with no increase in dynamic pressure while reducing the tunnel drive-power requirements. Studies have been made to determine the expected variations of Reynolds number and other parameters over wide ranges of Mach number, pressure, and temperature with due regard to avoiding liquefaction and adverse real-gas effects. Practical operational procedures have been developed in a low-speed prototype cryogenic wind tunnel. Aerodynamic experiments in the facility have demonstrated the theoretically predicted variations in Reynolds number and drive power. Force and moment measurements on a wing model mounted on a water-jacketed strain-gage sting balance have demonstrated the feasibility of operation of such balances in a cryogenic environment.
The development of cryogenic wind tunnels is reviewed with reference to the theory and advantages of cryogenic tunnels, problems common to wind tunnels and their solution, and application of cryogenic wind tunnels to high Reynolds number testing. It is shown that cryogenic wind tunnels can achieve full-scale Reynolds number with reasonable tunnel size, dynamic pressure, and drive power; the use of such tunnels also makes it possible to separate the effects of Reynolds number, Mach number, and aeroelasticity. Application of the cryogenic tunnel concept is illustrated by three examples, namely an atmospheric low-speed cryogenic tunnel, a 0.3-meter transonic cryogenic tunnel, and the National Transonic Facility now nearing completion.
Tests were conducted under a USAF/NASA Low Power Deicer program on two expulsive technologies to examine system performance on hardware representative of a modern aircraft part. The BF Goodrich Electro-Expulsive Deicing System and Pneumatic Impulse Ice Protection System were installed on a swept, compound curve, engine inlet component with varying leading edge radius, and tested through a range of icing and system operating conditions in the NASA Lewis Icing Research Tunnel. A description of the experimental procedure and results, including residual ice thickness, shed ice particle size, and changes in system energy/pressure characteristics are presented.
Tests were conducted under a USAF/NASA Low Power Deicer program on two expulsive technologies to examine system performance on hardware representative of a modern aircraft part. The BF Goodrich Electro-Expulsive Deicing System and Pneumatic Impulse Ice Protection system were installed on a swept, compound curve, engine inlet component with varying leading edge radius, and tested through a range of icing and system operating conditions in the NASA Lewis Icing Research Tunnel. A description of the experimental procedure and results, including residual ice thickness, shed ice particle size, and changes in system energy/pressure characteristics are presented.
Future Mars human lander missions will require using powered descent beginning at supersonic conditions, something which has never been done before at Mars. Significant aerosciences challenges exist due to interactions between the retrorocket exhaust plumes and the freestream flow that alter the aerodynamic behavior of the powered descent vehicle. Historically, wind tunnel tests have been used to study supersonic retropropulsion with inert gas exhaust simulants. Also, SpaceX has successfully decelerated the Falcon 9 first stage numerous times by using powered flight at supersonic conditions. On the computational side, Reynolds-Averaged Navier-Stokes flowfield simulations are regularly completed at full-scale conditions. However, the available ground and flight data do not provide a basis for calibrating the computational uncertainties for aerodynamic interference forces and moments on proposed Mars descent vehicles, either because of insufficient data or dissimilar vehicles geometries and/or conditions. Thus, additional ground testing is needed to continue addressing the aerosciences risks for large-scale human Mars landers. To that end, a retropropulsion wind tunnel test will be conducted in the NASA Langley Unitary Plan Wind Tunnel in 2020. The test is designed to improve upon similar past tests, both in terms of model design and measured data. The test campaign will use subscale model geometries derived from the two current NASA powered descent reference vehicles: a blunt low lift-to-drag vehicle and a more slender geometry. This paper covers the reference vehicles, test objectives, model design, scaling parameters, test matrix, and computational analysis to date.
The predicted upper and lower bounds power spectra for all of the cases and response items given in Volume 1 are plotted. The flight test power spectra are shown on each prediction plot for the nominal value of angle of attack that most closely agrees with the flexible angle for the prediction. The flight test and prediction conditions are given in tabular form for all cases considered.
A full-scale model of a supersonic STOVL single engine flighter aircraft employing an ejector to augment lift in hover and in low-speed flight was tested in the 40- by 80-ft and 80- by 120-ft test sections of the National Full-Scale Aerodynamics Complex located at the NASA Ames Research Center. The measured ejector augmentation ratio in hover met the design requirement of 1.6 and continued to provide the lift necessary in forward flight for good transition qualities. The up-and-away aerodynamics (ejector system stowed) were found to be conventional for this class of vehicle. The pitch control provided by the full-span blown flaps is sufficient to control the large pitching moments generated by ventral exhaust nozzle vectoring and propulsion induced aerodynamic effects such as the turning of the flow entrained into the ejectors.
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Abstract not provided.
The aerodynamic database for the Orion Launch Abort System (LAS) was developed largely from wind tunnel tests involving powered jet simulations of the rocket exhaust plumes, supported by computational fluid dynamics (CFD) simulations. The LAS contains three solid rocket motors used in various phases of an abort to provide propulsion, steering, and Launch Abort Tower (LAT) jettison from the Crew Module (CM). This paper describes a pair of wind tunnel experiments performed at transonic and supersonic speeds to determine the aerodynamic effects due to proximity and jet interactions during LAT jettison from the CM at the end of an abort. The tests were run using two different scale models at angles of attack from 150deg to 200deg , sideslip angles from -10deg to +10deg , and a range of powered thrust levels from the jettison motors to match various jet simulation parameters with flight values. Separation movements between the CM and LAT included axial and vertical translations as well as relative pitch angle between the two bodies. The paper details aspects of the model design, nozzle scaling methodology, instrumentation, testing procedures, and data reduction. Sample data are shown to highlight trends seen in the results.