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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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The ability of CFD simulations to serve as a surrogate for wind tunnel testing at high supersonic speeds has been evaluated for a sub-scale model of the Co-Optimization Blunt-body Re-entry Analysis-Mid-lift-to-drag Rigid Vehicle (CobraMRV) human Mars entry vehicle concept. The vehicle was tested at the Unitary Plan Wind Tunnel (UPWT) at the NASA Langley Research Center under flow conditions and surface control configurations relevant to the entry stage of a flight mission. The CFD simulations were performed prior to gaining access to test results in order to assess how blind predictions obtained using best practices compare to experiments. Solutions of empty tunnel simulations were used as inflow boundary condition for the CobraMRV simulations in a truncated portion of the test section. Different solvers and turbulence models were used by separate teams to assess sensitivity to numerical methods, physics, and users. After release of the test results, the pre-test computations were compared to the experimental results, and additional analyses have been conducted to explain observed discrepancies. The amount of time and resources dedicated to each phase of the computational work was logged for comparison to that required for wind tunnel tests, and to inform planning of future CFD data base development projects.
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An investigation was conducted in a transonic cryogenic tunnel to determine the effect of varying Reynolds number on the boattail drag of several wing-body configurations. This study was made at 0 deg angle of attack at Mach numbers from 0.6 to 0.9 for Reynolds numbers up to 67 x 1 million (based on distance from the nose to the start of the boattail). Results indicate that as the Reynolds number was increased the boattail static pressure coefficients in the expansion region of the boattail became more negative while those in the recompression region became more positive. Results show that there was only a small effect of Reynolds number of boattail pressure drag.
An experimental investigation was conducted in a water tunnel to identify the effects of apex blowing on two delta-wing models undergoing constant pitch-rate motion. One wing was of 60-degree sweep and the other was of 76-degree sweep. Flow visualization methods were utilized to determine vortex burst locations for a wide range of pitch-up (nose up) and plunge (nose down) rates, apex jet strengths, and blowing directions. Results indicate that blowing along the 60-degree wing vortex core or along the 76-degree wing centerline results in a notable improvement in vortex behavior under both static and dynamic conditions. The results are most dramatic during dynamic plunging (nose down) conditions, where blowing resulted in the reformation of vortices with significant length.
An investigation was made in the Langley low-turbulence pressure tunnel to determine the effect of size and location of a sandpaper type of roughness on the Reynolds number for transition. Transition was observed by means of a hot-wire anemometer located at various chordwise stations for each position of the roughness. These observations indicated that when the roughness is sufficiently submerged in the boundary layer to provide a substantially linear variation of boundary-layer velocity with distance from the surface up to the top of the roughness, turbulent "spots" begin to appear immediately behind the roughness when the Reynolds number based on the velocity at the top of the roughness height exceeds a value of approximately 600. At Reynolds numbers even slightly below the critical value (value for transition), the sandpaper type of roughness introduced no measurable disturbances into the laminar layer downstream of the roughness. The extent of the roughness area does not appear to have an important effect on the critical value of the roughness Reynolds number.
The trailing-wing sensor technique was used in the Langley V/STOL tunnel to determine the effectiveness of 11 combinations of the existing flight-spoiler segments on a jumbo-jet transport aircraft model when they were deflected as trailing-vortex-alleviation devices. All 11 of the flight-spoiler configurations investigated were effective in reducing the induced rolling moment on the trailing model. This investigation is an extension of earlier wind-tunnel and flight tests which showed that the existing flight spoilers on the jumbo-jet aircraft can be used as effective trailing-vortex-alleviation devices. Essentially, all of the reduction in induced rolling moment on the trailing-wing model was realized at a spoiler deflection of 45 deg for single-spoiler configurations, 30 for two-spoiler configurations, and 15 deg for both the three- and four-spoiler configurations. Of the 11 flight-spoiler configurations investigated, the most promising configuration for trailing-vortex abatement on the jumbo-jet aircraft appears to be the three inboard flight spoilers deflected 15 deg.
The Linear Aerospike SR-71 Experiment (LASRE) is presently being conducted to test a 20-percent-scale version of the Linear Aerospike rocket engine. This rocket engine has been chosen to power the X-33 Single Stage to Orbit Technology Demonstrator Vehicle. The rocket engine was integrated into a lifting body configuration and mounted to the upper surface of an SR-71 aircraft. This paper presents stability and control results and performance results from the envelope expansion flight tests of the LASRE configuration up to Mach 1.8 and compares the results with wind tunnel predictions. Longitudinal stability and elevator control effectiveness were well-predicted from wind tunnel tests. Zero-lift pitching moment was mispredicted transonically. Directional stability, dihedral stability, and rudder effectiveness were overpredicted. The SR-71 handling qualities were never significantly impacted as a result of the missed predictions. Performance results confirmed the large amount of wind-tunnel-predicted transonic drag for the LASRE configuration. This drag increase made the performance of the vehicle so poor that acceleration through transonic Mach numbers could not be achieved on a hot day without depleting the available fuel.
A wind tunnel experiment was conducted in the NASA Langley Research Center (LaRC) Unitary Plan Wind Tunnel (UPWT) to determine the effects of passive surface porosity on vortex flow interaction about a general research fighter configuration at supersonic speeds. Optical flow measurement and flow visualization techniques were used and included pressure-sensitive paint (PSP), schlieren, and laser vapor screen (LVS) These techniques were combined with force and moment and conventional electronically-scanned pressure (ESP) measurements to quantify and to visualize the effects of flow-through porosity applied to a wing leading-edge extension (LEX) mounted to a 65 deg cropped delta wing model.
A wind tunnel experiment was conducted in the NASA Langley Research Center (LaRC) Unitary Plan Wind Tunnel (UPW7) to determine the effects of passive surface porosity on vortex flow interactions about a general research fighter configuration at supersonic speeds. Optical flow measurement and flow visualization techniques were used and included pressure-sensitive paint (PSP), schlieren, and laser vapor screen (LVS). These techniques were combined with force and moment and conventional electronically-scanned pressure (ESP) measurements to quantify and to visualize the effects flow-through porosity applied to a wing leading-edge extension (LEX) mounted to a 65 deg cropped delta wing model.
Wind tunnel tests to predict the aircraft wake turbulence due to the tip trailing vortex are discussed. A yawhead pressure probe was used in a subsonic wind tunnel to obtain detailed mean flow measurements at stations up to 30 chordlengths downstream in an aircraft trailing vortex. Mass injection at the wingtip was shown to hasten the decay of the trailing vortex. A theoretical method is presented to show the effect which the circulation distribution on the wing has on the structure of the outer portion of the vortex.
The aerodynamic characteristics of a typical modern general aviation airfoil were investigated with and without a pneumatic boot ice protection system. The ice protection effectiveness of the boot was studied. This includes the change in drag on the airfoil with the boot inflated and deflated, the change in drag due to primary and residual ice formation, drag change due to cumulative residual ice formation, and parameters affecting boot effectiveness. Boot performance was not affected by tunnel total temperature or velocity. Marginal effect in performance was associated with angle of attack. Significant effects on performance were caused by variations in droplet size, LWC, ice cap thickness inflation pressure, and surface treatment.
Inelastic electron tunneling spectroscopy is a useful technique for the study of vibrational modes of molecules adsorbed on the surface of oxide layers in a metal-insulator-metal tunnel junction. The technique involves studying the effects of adsorbed molecules on the tunneling spectrum of such junctions. The data give useful information about the structure, bonding, and orientation of adsorbed molecules. One of the major advantages of inelastic electron tunneling spectroscopy is its sensitivity. It is capable of detecting on the order of 10 to the 10th molecules (a fraction of a monolayer) on a 1 sq mm junction. It has been successfully used in studies of catalysis, biology, trace impurity detection, and electronic excitations. Because of its high sensitivity, this technique shows great promise in the area of solid-state electronic chemical sensing.
An investigation has been conducted in the static test facility of the Langley 16-Foot Transonic Tunnel to determine the effects of rounding and rotating the port corner of a nonaxisymmetric thrust reversing port. High pressure air was used to simulate jet exhaust at nozzle pressure ratios up to 6.0. The investigation indicates that using a rounded corner in the reverser port passage increases the discharge coefficient significantly compared to that of a sharp cornered thrust reversing port. Incorporating a rotating cylinder into the rounded corner reverser port passage provides no additional increase at cylinder revolutions per minute up to 10,000. (Revolutions per minute of 390,000 were necessary to achieve a surface velocity equal to the duct velocity.) Highest levels of discharge coefficient are obtained by providing a large, round, cylindrical surface in the port upstream wall. Locating the reverser port in a constant area duct results in higher discharge coefficients than locating the port in a convergent duct.
Wind tunnel interference corrections have direct impact on measured propeller efficiency. A systematic series of wind tunnel tests was done in the porous-wall NASA Lewis 8- by 6-Foot Wind Tunnel to determine the wind tunnel interference corrections to the NASA Lewis counterrotation propeller test data. The test results were compared with calculations from a potential flow code to determine the interference corrections. At a Mach number of 0.8, the interference corrections resulted in a -0.008 Mach number correction which reduced the counterrotation propeller net efficiency data by 0.46 percent at the reduced Mach number. Additional wind tunnel tests were done to measure the effect of propeller thrust on wind tunnel wall interference. No wall interference corrections due to propeller thrust were found necessary for the high speed counterrotation propeller data obtained in the porous wall NASA Lewis 8- by 6-Foot Wind Tunnel.