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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 145 records · Page 8
All sourcesNASA NTRSDOE OSTI
NASA NTRS2016

Assessment of JVX Proprotor Performance Data in Hover and Airplane-Mode Flight Conditions

A 0.656-scale V-22 proprotor, the Joint Vertical Experimental (JVX) rotor, was tested at the NASA Ames Research Center in both hover and airplane-mode (high-speed axial flow) flight conditions, up to an advance ratio of 0.562 (231 knots). This paper examines the two principal data sets generated by those tests, and includes investigations of hub spinner tares, torque/thrust measurement interactions, tunnel blockage effects, and other phenomena suspected of causing erroneous measurements or predictions. Uncertainties in hover and high-speed data are characterized. The results are reported here to provide guidance for future wind tunnel tests, data processing, and data analysis.

Proprotor↗
NASA NTRS2020

Full Aerothermal Characterization of the 6- by 9-Foot Icing Research Tunnel (2019 Test)

A full aerothermal characterization of the NASA Glenn Icing Research Tunnel (IRT) was completed in August 2019. A mandatory maintenance period required the test entry to be separated into two phases: the first occurring in April 2019 and the second in August 2019. The hardware used during this test entry consisted of the 6-foot quick-check rake, 6-foot traversing rake, RTD array, and hot-wire rake. The 6-foot quick check rake collects static pressure, total pressure, total temperature, and flow angularity data along the centerline of the test section. The data from the 6-foot quick-check rake were used to generate test section calibration relationships for total and static pressure as well as implement statistical process control. The 6-foot traversing rake collects flow uniformity data across the test section for various flow parameters, such as static pressure, total pressure, and flow angularity. Total temperature flow uniformity data is mapped by the RTD array and is used to generate the test section calibration relationship for total temperature. The hot-wire rake measures turbulence intensity at three lateral locations and characterizes airspeed unsteadiness in the test section. The test entry resulted in updated calibration relationships and a full characterization of test section flow quality and flow uniformity in the IRT.

Wind↗
NASA NTRS2013

Acoustic Characterization of Compact Jet Engine Simulator Units

Two dual-stream, heated jet, Compact Jet Engine Simulator (CJES) units are designed for wind tunnel acoustic experiments involving a Hybrid Wing Body (HWB) vehicle. The newly fabricated CJES units are characterized with a series of acoustic and flowfield investigations to ensure successful operation with minimal rig noise. To limit simulator size, consistent with a 5.8% HWB model, the CJES units adapt Ultra Compact Combustor (UCC) technology developed at the Air Force Research Laboratory. Stable and controllable operation of the combustor is demonstrated using passive swirl air injection and backpressuring of the combustion chamber. Combustion instability tones are eliminated using nonuniform flow conditioners in conjunction with upstream screens. Through proper flow conditioning, rig noise is reduced by more than 20 dB over a broad spectral range, but it is not completely eliminated at high frequencies. The low-noise chevron nozzle concept designed for the HWB test shows expected acoustic benefits when installed on the CJES unit, and consistency between CJES units is shown to be within 0.5 dB OASPL.

Doty, Michael J.↗
NASA NTRS1976

Simulation of flight-type engine fan noise in the NASA-Lewis 9 x 15 anechoic wind tunnel

A major problem in the measurement of aircraft engine fan noise is the difficulty of simulating, in a ground-based facility, the noise that occurs during flight. Flight-type noise as contrasted to the usual ground-static test noise exhibits substantial reductions in both (1) the time unsteadiness of tone noise and (2) the mean level of tones calculated to be nonpropagating or cut-off. A model fan designed with cut-off of the fundamental tone was acoustically tested in the anechoic wind tunnel under both static and tunnel flow conditions. The properties that characterize flight-type noise were progressively simulated with increasing tunnel flow. The distinctly lobed directivity pattern of propagating rotor/stator interaction modes was also observed. The results imply that the excess noise attributed to the ingestion of the flow disturbances that prevail near most static test facilities was substantially reduced with tunnel flow. The anechoic wind tunnel appears to be a useful facility for applied research on aircraft engine fan noise under conditions of simulated flight.

Heidmann, M. F.↗
NASA NTRS2018

Measurements of Freestream Fluctuations in the NASA Langley 20-Inch Mach 6 Tunnel

An experimental campaign was conducted to measure and to characterize the freestream disturbance levels in the NASA Langley Research Center 20-Inch Mach 6 Wind Tunnel. A pitot rake was instrumented with fast pressure transducers, hot wires, and an atomic layer thermopile to quantify the fluctuation levels of pressure, mass flux, and heat flux, respectively. In conjunction with these probe-based measurements, focused laser differential interferometry was used to optically measure density fluctuations. Measurements were made at five nominal different unit Reynolds numbers ranging from (3.28 to 26.5) times 10 (sup 6) per meter. The rake was positioned at two different stream-wise locations and several different roll angles to measure flow uniformity within the test section. In general, noise levels were spatially consistent within the tested region. Pitot pressure fluctuation levels ranged from 0.84 percent at the highest Reynolds number tested to 1.89 percent at the lowest Reynolds number tested. Freestream mass-flux fluctuations remained relatively constant between 1.8-2.5 percent of the freestream. The pressure transducers were also used to determine the dominant disturbance speed and angle of propagation. The disturbances were estimated to travel at approximately 54-81 percent of the freestream speed at an angle of approximately 21-44 degrees from the freestream direction, but these measurements had a significant amount of uncertainty. A comparison to previous measurements of pressure made in 2012 and of mass flux made in 1994 show almost no change in the RMS (Root Mean Square) fluctuation of these flow quantities.

Chou, Amanda↗
NASA NTRSDate not supplied

Overview of the Integrated Adaptive Wing Technology Maturation Wind-Tunnel Test Objectives

The Integrated Adaptive Wing Technology Maturation wind-tunnel test is the culmination of a cooperative effort between NASA and Boeing to develop control techniques that can improve the performance and safety of higher aspect ratio flexible transport aircraft. The wind-tunnel test article, to be tested at the NASA Langley Research Center Transonic Dynamics Tunnel, is a semispan model of a modern transport configuration with an aspect ratio 13.5 wing based on the NASA Common Research Model. Multiobjective control laws have been developed targeting drag optimization, maneuver load alleviation, gust load alleviation, and modal suppression. In this paper, the integrated model design, active control system, and sensor suite will be presented, and pretest modeling and characterization activities will be described. A summary of the planned wind-tunnel test entries will provide an overview of the experimental approach and expected outcomes.

wind-tunnel↗
NASA NTRSDate not supplied

Overview of the Integrated Adaptive Wing Technology Maturation Wind-Tunnel Test Objectives

The Integrated Adaptive Wing Technology Maturation wind-tunnel test is the culmination of a cooperative effort between NASA and Boeing to develop control techniques that can improve the performance and safety of higher aspect ratio flexible transport aircraft. The wind-tunnel test article, to be tested at the NASA Langley Research Center Transonic Dynamics Tunnel, is a semispan model of a modern transport configuration based on the NASA Common Research Model with an aspect ratio 13.5 wing. Multiobjective control laws have been developed targeting drag optimization, maneuver load alleviation, gust load alleviation, and active flutter suppression. In this paper, the integrated model design, active control system, and sensor suite will be presented, and pretest modeling and characterization activities will be described. A summary of the planned wind-tunnel test entries will provide an overview of the experimental approach and expected outcomes.

wind-tunnel↗
NASA NTRS1985

Background noise measurements from jet exit vanes designed to reduced flow pulsations in an open-jet wind tunnel

Many open jet wind tunnels experience pulsations of the flow which are typically characterized by periodic low frequency velocity and pressure variations. One method of reducing these fluctuations is to install vanes around the perimeter of the jet exit to protrude into the flow. Although these vanes were shown to be effective in reducing the fluctuation content, they can also increase the test section background noise level. The results of an experimental acoustic program in the Langley 4- by 7-Meter Tunnel is presented which evaluates the effect on tunnel background noise of such modifications to the jet exit nozzle. Noise levels for the baseline tunnel configuration are compared with those for three jet exit nozzle modifications, including an enhanced noise reduction configuration that minimizes the effect of the vanes on the background noise. Although the noise levels for this modified vane configuration were comparable to baseline tunnel background noise levels in this facility, installation of these modified vanes in an acoustic tunnel may be of concern because the noise levels for the vanes could be well above background noise levels in a quiet facility.

Hoad, D. R.↗
NASA NTRS2016

Computational Evaluation of Inlet Distortion on an Ejector Powered Hybrid Wing Body at Takeoff and Landing Conditions

Due to the aft, upper surface engine location on the Hybrid Wing Body (HWB) planform, there is potential to shed vorticity and separated wakes into the engine when the vehicle is operated at off-design conditions and corners of the envelope required for engine and airplane certification. CFD simulations were performed of the full-scale reference propulsion system, operating at a range of inlet flow rates, flight speeds, altitudes, angles of attack, and angles of sideslip to identify the conditions which produce the largest distortion and lowest pressure recovery. Pretest CFD was performed by NASA and Boeing, using multiple CFD codes, with various turbulence models. These data were used to make decisions regarding model integration, characterize inlet flow distortion patterns, and help define the wind tunnel test matrix. CFD was also performed post-test; when compared with test data, it was possible to make comparisons between measured model-scale and predicted full-scale distortion levels. This paper summarizes these CFD analyses.

Wind Tunnel Testing↗
NASA NTRSDate not supplied

CFD Predictions of Fluctuating Pressure Environments on NASA’s SLS Ascent Unsteady Aerodynamics Wind Tunnel Test

This paper details work that is being conducted on predicting fluctuating pressure environments on NASA launch vehicles using computational fluid dynamics (CFD). The accurate characterization of these aeroacoustics environments is necessary in the analysis of vehicle structural health and aerodynamic performance and in the vehicle design process. High-fidelity simulations of the unsteady flow over NASA’s SLS Ascent Unsteady Aerodynamics wind tunnel Test (AUAT) using the hybrid Reynolds-averaged Navier-Stokes (RANS)-large eddy simulations (LES) methodologies in Mississippi States’ Loci/CHEM solver are presented. The two available methods, Nelson-Nichols and dynamic hybrid RANS-LES (DHRL), are compared head-to-head on identical grids at two Mach numbers. At the high-subsonic Mach number, locally supersonic flow expanding over a shoulder induces a separation-reattachment system that is predicted poorly by the Nelson-Nichols method due to an under-prediction in unsteady content. The DHRL solution predicts the spatial extent and frequency domain response of this system well but under-predicts the peak noise levels due to a delay in the RANS-to-LES transition. At the low-supersonic Mach number, compression corner dynamics dominate the unsteady flow. While the Nelson Nichols solution captures this phenomenon well, another solution on a refined grid indicates that some model stress depletion is occurring. The DHRL solution again performs well at this Mach number and matches the wind tunnel data fairly closely, but perhaps most impressively, is fairly agnostic to increasing grid resolution. The findings in this paper are expected to be generally applicable and will be used to guide the prediction of aeroacoustics environments of other NASA launch vehicles.

CFD↗
NASA NTRS1990

Advanced instrumentation for aircraft icing research

A compact and rugged probe based on the phase Doppler method was evaluated as a means for characterizing icing clouds using airborne platforms and for advancing aircraft icing research in large scale wind tunnels. The Phase Doppler Particle Analyzer (PDPA) upon which the new probe was based is now widely recognized as an accurate method for the complete characterization of sprays. The prototype fiber optic-based probe was evaluated in simulated aircraft icing clouds and found to have the qualities essential to providing information that will advance aircraft icing research. Measurement comparisons of the size and velocity distributions made with the standard PDPA and the fiber optic probe were in excellent agreement as were the measurements of number density and liquid water content. Preliminary testing in the NASA Lewis Icing Research Tunnel (IRT) produced reasonable results but revealed some problems with vibration and signal quality at high speeds. The cause of these problems were identified and design changes were proposed to eliminate the shortcomings of the probe.

Bachalo, W.↗
NASA NTRS2005

Development and Characterization Testing of an Air Pulsation Valve for a Pulse Detonation Engine Supersonic Parametric Inlet Test Section

In pulse detonation engines, the potential exists for gas pulses from the combustor to travel upstream and adversely affect the inlet performance of the engine. In order to determine the effect of these high frequency pulses on the inlet performance, an air pulsation valve was developed to provide air pulses downstream of a supersonic parametric inlet test section. The purpose of this report is to document the design and characterization tests that were performed on a pulsation valve that was tested at the NASA Glenn Research Center 1x1 Supersonic Wind Tunnel (SWT) test facility. The high air flow pulsation valve design philosophy and analyses performed are discussed and characterization test results are presented. The pulsation valve model was devised based on the concept of using a free spinning ball valve driven from a variable speed electric motor to generate air flow pulses at preset frequencies. In order to deliver the proper flow rate, the flow port was contoured to maximize flow rate and minimize pressure drop. To obtain sharp pressure spikes the valve flow port was designed to be as narrow as possible to minimize port dwell time.

Tornabene, Robert↗
NASA NTRS2018

Regression Analysis and Calibration Recommendations for the Characterization of Balance Temperature Effects

Analysis and use of temperature-dependent wind tunnel strain-gage balance calibration data are discussed in the paper. First, three different methods are presented and compared that may be used to process temperature-dependent strain-gage balance data. The first method uses an extended set of independent variables in order to process the data and predict balance loads. The second method applies an extended load iteration equation during the analysis of balance calibration data. The third method uses temperature-dependent sensitivities for the data analysis. Physical interpretations of the most important temperature-dependent regression model terms are provided that relate temperature compensation imperfections and the temperature-dependent nature of the gage factor to sets of regression model terms. Finally, balance calibration recommendations are listed so that temperature-dependent calibration data can be obtained and successfully processed using the reviewed analysis methods.

wind tunnel test↗
NASA NTRS1982

Water tunnel flow visualization and wind tunnel data analysis of the F/A-18

Six degree of freedom studies were utilized to extract a band of yawing and rolling moment coefficients from the F/A-18 aircraft flight records. These were compared with 0.06 scale model data obtained in a 16T wind tunnel facility. The results, indicate the flight test yawing moment data exhibit an improvement over the wind tunnel data to near neutral stability and a significant reduction in lateral stability (again to anear neutral level). These data are consistent with the flight test results since the motion was characterized by a relatively slo departure. Flight tests repeated the slow yaw departure at M 0.3. Only 0.16 scale model wind tunnel data showed levels of lateral stability similar to the flight test results. Accordingly, geometric modifications were investigated on the 0.16 scale model in the 30x60 foot wind tunnel to improve high angle of attack lateral stability.

Erickson, G. E.↗
NASA NTRS1976

Simulation of Flight-Type Engine Fan Noise in the NASA-Lewis 9X15 Anechoic Wind Tunnel

Flight type noise as contrasted to the usual ground static test noise exhibits substantial reductions in the time unsteadiness of tone noise, and in the mean level of tones calculated to be nonpropagating or cut-off. A model fan designed with cuttoff of the fundamental tone was acoustically tested in the anechoic wind tunnel under both static and tunnel flow conditions. The properties that characterize flight type noise were progressively simulated with increasing tunnel flow. The distinctly lobed directivity pattern of propagating rotor/stator interaction modes was also observed. Excess noise attributed to the ingestion of the flow disturbances that prevail near most static test facilities is substantially reduced with tunnel flow.

Heidmann, M. F.↗
NASA NTRS2016

Active Flow Control Using Sweeping Jet Actuators on a Semi-Span Wing Model

Wind tunnel experiments were performed using active flow control on an unswept semispan wing model with a 30% chord trailing edge flap to aid in the selection of actuators for a planned high Reynolds number experiment. Two sweeping jet actuator sizes were investigated to determine the influence of actuator size on the active flow control system efficiency. Sweeping jet actuators with orifice sizes of 1 mm x 2 mm and 2 mm x 4 mm were selected because of the differences in actuator jet sweep angle. The parameters that were varied include actuator momentum, freestream velocity, and trailing edge flap deflection angle. Steady and unsteady pressure data, Particle Image Velocimetry data, and force and moment data were acquired to assess the performance of the two actuators. In addition to the wind tunnel experiments, benchtop studies of the actuators were performed to characterize the jets produced by each actuator. Benchtop investigations of the smaller actuator reveal that the jet exiting the actuator has a reduced sweep angle compared to published data for larger versions of this type of actuator. The larger actuator produces an oscillating jet that attaches to the external di↵user walls at low supply pressures and produces the expected sweep angles. The AFC results using the smaller actuators show that while the actuators can control flow separation, the selected spacing of 3.3 cm may be too large due to the reduced sweep angle. In comparison, the spacing for the larger actuators, 6.6 cm, appears to be optimal for the Mach numbers investigated. Particle Image Velocimetry results are presented and show how the wall jets produced by the actuators cause the flow to attach to the flap surface.

Melton, LaTunia Pack↗
NASA NTRS2021

Unsteady Propagation and Mean Corrections in Open-Jet and Kevlar® Wind Tunnels

Two types of aeroacoustic wind-tunnel test section configurations have been tested in the NASA Langley Quiet Flow Facility. The first is a more traditional open-jet configuration, where test section flow passes unbounded through the facility anechoic chamber. The second is a Kevlar® panel configuration, where a tensioned Kevlar sheet constrains the test section flow from the facility anechoic chamber. For both configurations, acoustic instrumentation is in the surrounding quiescent space. Both configurations are evaluated with a laser-based pulsed acoustic source, which provides unique capability for assessing the facility unsteady acoustic propagation characteristics. Metrics based on the wander and spread of the pulses are evaluated, and show that measurements using Kevlar walls experience dramatically reduced unsteady effects when compared to the open-jet configuration. This leads to a corresponding improvement in coherence between microphones with the Kevlar configuration, by reducing the variation in magnitude and phase differences between channels. Magnitude corrections for propagation through Kevlar as compared to open-jet propagation are calculated. Although limitations in the experimental setup make quantitative analysis difficult, qualitative analysis shows Kevlar magnitude corrections similar to those determined in previous literature. Directivity effects beyond those already present for open-jet configurations are minimal. When compared with the open-jet configuration, the background noise produced by the Kevlar is found to be its one drawback, showing significantly greater levels at high frequencies.

wind tunnel testing, hybrid acoustic tunnel, facil↗
NASA NTRS1994

Experimental studies of transonic flow field near a longitudinally slotted wind tunnel wall

The results of detailed parametric experiments are presented for the near-wall flow field of a longitudinally slotted transonic wind tunnel. Existing data are reevaluated and new data obtained in the Langley 6- by 19-inch Transonic Wind Tunnel are presented and analyzed. In the experiments, researchers systematically investigate many pertinent wall-geometry variables such as the wall openness and the number of slots along with the free stream Mach number and model angle of attack. Flow field surveys on the plane passing through the centerline of the slot were conducted and are presented. The effects of viscosity on the slot flow are considered in the analysis. The present experiments, combined with those of previous investigations, give a more complete physical characterization of the flow near and through the slotted wall of a transonic wind tunnel.

Everhart, Joel L.↗