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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 469 records · Page 26

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NASA NTRS1989

Cruise noise of the SR-2 propeller model in a wind tunnel

Noise data on the SR-2 model propeller were taken in the NASA Lewis Research Center 8- by 6-Foot Wind Tunnel. The maximum blade passing tone rises with increasing helical tip Mach number to a peak level at a helical tip Mach number of about 1.05; then it remains the same or decreases at higher helical tip Mach numbers. This behavior, which has been observed with other propeller models, points to the possibility of using higher propeller tip speeds to limit airplane cabin noise while maintaining high flight speed and efficiency. Noise comparisons of the straight-blade SR-2 propeller and the swept-blade SR-7A propeller showed that the tailored sweep of the SR-7A appears to be the cause of both lower peak noise levels and a slower noise increase with increasing helical tip Mach number.

Dittmar, James H.↗

NASA NTRS2010

Incorporation of SemiSpan SuperSonic Transport (S4T) Aeroservoelastic Models into SAREC-ASV Simulation

The Simulink-based Simulation Architecture for Evaluating Controls for Aerospace Vehicles (SAREC-ASV) was modified to incorporate linear models representing aeroservoelastic characteristics of the SemiSpan SuperSonic Transport (S4T) wind-tunnel model. The S4T planform is for a Technology Concept Aircraft (TCA) design from the 1990s. The model has three control surfaces and is instrumented with accelerometers and strain gauges. Control laws developed for wind-tunnel testing for Ride Quality Enhancement, Gust Load Alleviation, and Flutter Suppression System functions were implemented in the simulation. The simulation models open- and closed-loop response to turbulence and to control excitation. It provides time histories for closed-loop stable conditions above the open-loop flutter boundary. The simulation is useful for assessing the potential impact of closed-loop control rate and position saturation. It also provides a means to assess fidelity of system identification procedures by providing time histories for a known plant model, with and without unmeasured turbulence as a disturbance. Sets of linear models representing different Mach number and dynamic pressure conditions were implemented as MATLAB Linear Time Invariant (LTI) objects. Configuration changes were implemented by selecting which LTI object to use in a Simulink template block. A limited comparison of simulation versus wind-tunnel results is shown.

Christhilf, David M.↗

NASA NTRS1989

Analysis of an unswept propfan blade with a semiempirical dynamic stall model

The time history response of a propfan wind tunnel model with dynamic stall is studied analytically. The response obtained from the analysis is compared with available experimental data. The governing equations of motion are formulated in terms of blade normal modes which are calculated using the COSMIC-NASTRAN computer code. The response analysis considered the blade plunging and pitching motions. The lift, drag and moment coefficients for angles of attack below the static stall angle are obtained from a quasi-steady theory. For angles above static stall angles, a semiempirical dynamic stall model based on a correction to angle of attack is used to obtain lift, drag and moment coefficients. Using these coefficients, the aerodynamic forces are calculated at a selected number of strips, and integrated to obtain the total generalized forces. The combined momentum-blade element theory is used to calculate the induced velocity. The semiempirical stall model predicted a limit cycle oscillation near the setting angle at which large vibratory stresses were observed in an experiment. The predicted mode and frequency of oscillation also agreed with those measured in the experiment near the setting angle.

Reddy, T. S. R.↗

NASA NTRSDate not supplied

Launch Vehicle Loads Analysis Using Pressure-Sensitive Paint

Pressure transducers have been the instrumentation of choice for measuring unsteady flow phenomena. With recent advances in high-speed cameras, high-powered LEDs, and fast-response, pressure-sensitive paint, the unsteady pressure-sensitive paint (uPSP) technique has become a valuable alternative for production wind tunnel facilities, enabling time-resolved measurements of unsteady pressure fluctuations over a dense spatial grid on a wind tunnel model. Launch vehicle ground tests have proven to be a particularly well-matched application for uPSP due to the high signal level relative to tunnel background acoustics, relatively simple camera optical access, and rigidity of the model in wind-on test conditions. This presentation will highlight recent advances in data reduction of uPSP measurement data from recent launch vehicle wind tunnel tests at the NASA Ames Unitary Plan Wind Tunnel Complex (UPWT). The system can provide both localized surface pressure spectra as well as regional or zonal estimates of turbulence correlation model parameters. In addition, integrated vehicle-scale loads can be provided for buffet analysis. Data is reduced at the on-premise NASA Advanced Supercomputer (NAS) Division for just-in-time delivery of results during an ongoing wind tunnel test.

Pressure-Sensitive Paint↗

NASA NTRSDate not supplied

Launch Vehicle Loads Analysis Using Pressure-Sensitive Paint

pressure-sensitive paint↗

NASA NTRS1982

In-flight propeller flow visualization using fluorescent minitufts

Extension of fluorescent minutuft method to in-flight flow visualization on propellers is described. Extremely thin nylon monofilament for the minitufts, is used in a process of attaching them to the test surface with small drops of lacquer-like adhesive, and the use of fluorescence photography for recording the minituft patterns. Using this method, thousands of minitufts can be applied to small, high speed wind tunnel models without affecting the airflow. The minitufts can remain in place throughout a wind tunnel test, permitting nonintrusive flow visualization data to be acquired at any time.

Crowder, J. P.↗

NASA NTRS1987

The design of a joined wing flight demonstrator aircraft

A joined-wing flight demonstrator aircraft has been developed at the NASA Ames Research Center in collaboration with ACA Industries. The aircraft is designed to utilize the fuselage, engines, and undercarriage of the existing NASA AD-1 flight demonstrator aircraft. The design objectives, methods, constraints, and the resulting aircraft design, called the JW-1, are presented. A wind-tunnel model of the JW-1 was tested in the NASA Ames 12-foot wind tunnel. The test results indicate that the JW-1 has satisfactory flying qualities for a flight demonstrator aircraft. Good agreement of test results with design predictions confirmed the validity of the design methods used for application to joined-wing configurations.

Smith, S. C.↗

NASA NTRSDate not supplied

Noise Reduction and Calibration for Aeroacoustic Loads from Unsteady Pressure-Sensitive Paint

Unsteady pressure-sensitive paint (uPSP) is a significant new technology for high spatial resolution measurement of complex dynamic surface pressure loading on a wind tunnel model, to better define flight vehicle buffet and aeroacoustic loads. Compared with the very sparse surface pressure measurements of traditional sensors, the pressure sensitivity of uPSP light reflectivity provides a visual measure of surface pressure fluctuations over a wide area illuminated by high intensity light, captured by a high-speed digital camera. Panda et al (AIAA Journal, May 2019) have shown good data is recoverable up to 10 kHz. In cases where available uPSP luminescence is low for a painted wind tunnel test model, significant noise levels have been observed primarily due to the discrete photon arrival variability of low intensity light (termed “shot noise”). In this case, uPSP measurements can exhibit unusually low signal to noise ratio (SNR of 0.1 or less). This paper will report on latest work to recover calibrated uPSP measurements which are noise free over a useful dynamic range of 10-15 dB. It will be shown how the distinct spatial correlations of both the shot noise and the underlying aeroacoustic surface pressure play an important role in obtaining a reliable measurement.

pressure-sensitive paint↗

NASA NTRS2014

Pressure Sensitive Paint Applied to Flexible Models Project

One gap in current pressure-measurement technology is a high-spatial-resolution method for accurately measuring pressures on spatially and temporally varying wind-tunnel models such as Inflatable Aerodynamic Decelerators (IADs), parachutes, and sails. Conventional pressure taps only provide sparse measurements at discrete points and are difficult to integrate with the model structure without altering structural properties. Pressure Sensitive Paint (PSP) provides pressure measurements with high spatial resolution, but its use has been limited to rigid or semi-rigid models. Extending the use of PSP from rigid surfaces to flexible surfaces would allow direct, high-spatial-resolution measurements of the unsteady surface pressure distribution. Once developed, this new capability will be combined with existing stereo photogrammetry methods to simultaneously measure the shape of a dynamically deforming model in a wind tunnel. Presented here are the results and methodology for using PSP on flexible surfaces.

model deformation↗

NASA NTRS1983

Approach to establishing the effect of aeroelasticity on aerodynamic characteristics of the space shuttle orbiter

The static aeroelastic effects on the longitudinal stability and elevon/aileron effectiveness of the space transportation system (STS) Space Shuttle orbiter were estimated by a simplified approach called the elevon torsional stiffness (ETS) method. This method employs rigid model wind tunnel test results to predict aeroelastic effects. Lateral/directional stability and rudder effectiveness were based on results of a wind tunnel test in which a flexible tail model was used. Comparisons with selective flight data are made in this paper. Results of correlations with flight data (although limited at the present time) verify the predicted aeroelastic effects for the orbiter. The orbiter's structural characteristics are such that the effects of aeroelasticity, whether estimated using analytical techniques or simplified methods, do not appear to affect the vehicle performance to any great extent. The large amount of scatter in the flight-extracted data made verification of the aeroelastic corrections very difficult. Generally, the simplified elevon torsional stiffness method provided better correlation with flight test results than he analytical method and reduced the verification effort and cost.

Schlosser, D. C.↗

NASA NTRS1988

The composite blade structural analyzer (COBSTRAN)

The use and application of the COBSTRAN (COmposite Blade STRuctural ANalyzer) computer code is presented. COBSTRAN was developed at NASA-Lewis and is currently being used for the design and analysis of aircraft engine ducted and unducted fan blades. The features of COBSTRAN are demonstrated for the modeling and analysis of a scaled down wind tunnel model propfan blade made from fiber composites. Comparison of analytical and experimental mode shapes and frequencies are shown, verifying the model development and analysis techniques used. The methodologies and programs developed for this analysis are directly applicable to other propfan blades.

Aiello, Robert A.↗

NASA NTRS1988

Theoretical evaluation of engine auxiliary inlet design for supersonic V/STOL aircraft

A higher order panel method is used to evaluate the potential flow of a two dimensional supersonic V/STOL inlet. A non-symmetric analytical inlet model is developed to closely match a wind tunnel model. The analytical inlet is analyzed for flow characteristics around the lower cowl lip and auxiliary inlets. The results are obtained from the output of a computer program that is based on the Hess Panel Method which determines source strengths of panels distributed over a three dimensional body. The analytical model was designed for the implementation of drooped/translated cowl lip and auxiliary inlets as flow improvement concepts. A 40 or 70 degree droop lip can be incorporated on the inlet to determine if these geometry modifications result in flow improvements which may reduce the propensity for flow separation on the interior portion of the lip. Auxiliary inlets are employed to decrease the mass flow over the inlet lip. Thus, the peak flow velocity is reduced at the lip which also lessens the likelihood of flow separation on the interior portion of the lip. A 2, 4, and 6 inch translated lip can be employed to also decrease mass flow over the inlet lower lip in the same manner as the auxiliary inlet. The performance results of the flow improvement concepts show that three possible inlet configurations provide a situation where separation is less likely to occur. A 70 degree droop lip maintains flow conditions such that attached flow over the lower cowl lip may exist for the entire angle of attack range studied. A 0 degree droop and translated lip combination provides similar results for the angle of attack range. The third configuration consists of a 0 degree droop and auxiliary inlet combination. This configuration provides slightly less favorable results than the other two, but still allows for conditions favorable to attached flow within the inlet.

Boles, Michael A.↗

NASA NTRS1991

Theoretical evaluation of engine auxiliary inlet design for supersonic V/STOL aircraft

A higher order panel method is used to evaluate the potential flow of a 2-D supersonic V/STOL inlet. A nonsymmetric analytical inlet model is developed to closely match a wind tunnel model tested at NASA-Lewis. The analytical inlet used is analyzed for flow characteristics around the lower cowl lip and auxiliary inlets. The results for this analysis are obtained for the output of a computer program produced by the McDonnell Douglas Corp. This program is based on the Hess Panel Method which determines source strengths of panel distributed over a 3-D body. The analytical model was designed for the implementation of a drooped/translated cowl lip and auxiliary inlets as flow improvement concepts. A 40 or 70 deg droop lip can be incorporated on the inlet to determine if these geometry changes result in flow improvements which may reduce the propensity for flow separation on the interior portion of the lip. Auxiliary inlets are used to decrease the mass flow over the inlet lip; thus, the peak flow velocity is reduced at the lip which also lessens the likelihood of flow separation on the interior portion of the lip. A 2, 4, and 6 in. translated lip can be used to also decrease mass flow over the inlet lower lip in the same manner.

Boles, Michael A.↗

NASA NTRS1988

A study of aeroelastic stability for the model support system of the National Transonic Facility

Oscillations of wind-tunnel models have been observed during testing in the National Transonic Facility. These oscillations have been the subject of an extensive investigation. As a part of this effort, a study of the aeroelastic stability of the model support structure has been performed. This structure is mathematically modelled as a wing and conventional flutter analysis is performed. The math model implemented both experimentally and numerically obtained modal characteristics. A technique for illustrating the flutter boundary for wind tunnels is demonstrated. Results indicate that the classical flutter boundary is well above the operating envelope of the facility. However, the analysis indicates a damping-dependent instability is present which is inherent in the design. One possible modification in the design has been evaluated which eliminates the predicted instability.

Strganac, Thomas W.↗

NASA NTRS1983

Eliminating Wind Tunnel Flow Breakdown

Undesirable vortexes near floor in small wind tunnels suppressed by simple device that alters flow pattern there. Air is injected along floor and interacts with backflow from wind-tunnel model. Results in smoother, more correct air-flow and to more-reliable wind-tunnel data.

Hackett, J. E.↗

NASA NTRS1943

Jet-boundary corrections for reflection-plane models in rectangular wind tunnel

A detailed method for determining the jet-boundary corrections for reflection-plane models in rectangular wind tunnels is presented. The method includes the determination of the tunnel span local distribution and the derivation of equations for the corrections to the angle of attack, the lift and drag coefficients, and the pitching-, rolling-, yawing-, and hinge-moment coefficients. The principle effects of aerodynamic induction and of the boundary-induced curvature of the streamlines have been considered. An example is included to illustrate the method. Numerical values of the more important corrections for reflection-plane models in 7 by 10-foot closed wind tunnels are presented.

Swanson, Robert S↗

NASA NTRS1983

Rod-Wall Sound Shield for Wind Tunnels

Test model is shielded from turbulence radiated from wind tunnel walls. Shield overcomes problems caused by leading-edge configuration and leading edge angle of inclination of previous designs. Has successfully maintained a laminar rather than turbulent boundary layer.

Creel, T. R. J.↗

NASA NTRSDate not supplied

Mars Sample Return, Sample Retrieval Lander, Reaction Control System Jet Interaction Supersonic Wind Tunnel Test Overview with CFD Predictions

NASA's Mars Sample Return campaign will be launching several missions over the next decade that will work together to return rock samples from Mars back to Earth. The Sample Retrieval Lander (SRL) will deliver the Mars Ascent Vehicle and fetch rover to the surface of Mars in 2006. Rock samples collected by the Mars 2020 Perseverance rover, landing in early 2021, will be loaded on the the ascent vehicle to be launched into Mars orbit for retrieval by yet another spacecraft. The Sample Retrieval Lander will be a blunt entry capsule similar to past Mars entry vehicles like Mars Science Laboratory and Viking. The vehicle will fly a guided entry, using a small lift vector produced by a non-zero trim angle of attack to eliminate downrange and crossrange position errors at the point of parachute deploy. This energy and heading management is achieved with a reaction control system (RCS) that directs the bank angle of the vehicle and also minimizes unwanted capsule dynamics. The reaction control system and control design is based on the Mars Science Laboratory and Mars 2020 RCS systems. However, due to packaging constraints, the backshell of this new entry vehicle has a different geometry than those earlier designs. To certify the RCS system for flight the project must characterize the jet plume interactions with the capsule backshell that could impair or significantly augment the RCS control authority. This characterization will be done through a combination of computational fluid dynamics (CFD) analysis and wind tunnel test. Two candidate arrangements of the RCS jets have been identified for the SRL vehicle and are currently under evaluation before final selection. The aero/RCS plume interactions of these candidate configurations have been measured in a supersonic wind tunnel test in NASA Langley's Unitary Plan Wind Tunnel. The test was conducted in the fall of 2020 and data is currently being reduced. An overview of the candidate RCS configurations are presented here with an overview of the wind tunnel model design, jet scaling and scaled nozzle design, and the test matrix. Preliminary CFD runs are presented with an assessment of the predicted plumes and their interaction with the wake flow of the vehicle. The predicted effects of the model sting is provided as well. This high fidelity wind tunnel test is being conducted much earlier in the SRL project than would normally be done. The test was funded as part of a CFD evaluation task funded by NASA's Aerosciences Evaluation and Test Capabilities Project. The objective of the evaluation task was to compare the ability of CFD to predict complex flows with data that can be measured in the Langley Unitary Plan Wind Tunnel. RCS Jet interactions were selected as a type of complex flow that is important to NASA missions. In addition to providing useful data to the SRL project, there was added emphasis on quantifying the accuracy of the CFD predictions and wind tunnel test data. An overview of the uncertainty quantification methodologies for computational and experimental portions of this test is presented.

blunt body↗