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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 271 records · Page 15
All sourcesNASA NTRSDOE OSTI
NASA NTRS2018

Performance Calculations for a Boundary-Layer-Ingesting Fan Stage from Sparse Measurements

A test of the Boundary Layer Ingesting-Inlet / Distortion-Tolerant Fan was completed in NASA Glenn's 8-Foot by 6-Foot supersonic wind tunnel. Inlet and fan performance were measured by surveys using a set of rotating rake arrays upstream and downstream of the fan stage. Surveys were conducted along the 100 percent speed line and a constant exit corrected flow line passing through the aerodynamic design point. These surveys represented only a small fraction of the data collected during the test. For other operating points, data was recorded as snapshots without rotating the rakes which resulted in a sparser set of recorded data. This paper will discuss analysis of these additional, lower measurement density data points to expand our coverage of the fan map. Several techniques will be used to supplement the snapshot data at test conditions where survey data also exists. The supplemented snapshot data will be compared with survey results to assess the quality of the approach. Effective methods will be used to analyze the data set for which only snapshots exist.

fan efficiency↗
NASA NTRS1957

Total-Pressure Distortion and Recovery of Supersonic Nose Inlet with Conical Centerbody in Subsonic Icing Conditions

Ice was formed on a full-scale unheated supersonic nose inlet in the NACA Lewis icing tunnel to determine its effect on compressor-face total-pressure distortion and recovery.Inlet angle of attack was varied from 0degrees to 12 degrees, free-stream Mach number from 0.17 to 0.28, and compressor-face Mach number from 0.10 to 0.47. Icing-cloud liquid-water content was varied from 0.65 to 1.8 grams per cubic meter at free-stream static air temperatures of 15 degrees and 0 degrees F. The addition of ice to the inlet components increased total-pressure-distortion levels and decreased recovery values compared withclear0air results, the losses increasing with time in ice. The combination of glaze ice, high corrected weight flow, and high angle of attack yielded the highest levels of distortion and lowest values of recovery. The general character of compressor-face distortion with an iced inlet was the same as that for the clean inlet, the total-pressure gradients being predominantly radial, with circumferential gradients occurring at angle of attack. At zero angle of attack, free-stream Mach number of 0.27, and a constant corrected weight flow of 150 pounds per second (compressor-face Mach number of 0.43), compressor-face total-pressure-distortion level increased from about 6 percent in clear air to 12 percent after 21 minutes of heavy glaze icing; concurrently, total-pressure recovery decreased from about 0.98 to 0.945. For the same operating conditions but with the inlet at 12 deg angle of attack, a change in distortion level occurred from about 9 percent in clear air to 14 percent after 2-1/4 minutes of icing, with a decrease in recovery from about 0.97 to 0.94.

Gelder, Thomas F↗
NASA NTRS2010

Boundary-layer-ingesting inlet flow control system

A system for reducing distortion at the aerodynamic interface plane of a boundary-layer-ingesting inlet using a combination of active and passive flow control devices is disclosed. Active flow control jets and vortex generating vanes are used in combination to reduce distortion across a range of inlet operating conditions. Together, the vortex generating vanes can reduce most of the inlet distortion and the active flow control jets can be used at a significantly reduced control jet mass flow rate to make sure the inlet distortion stays low as the inlet mass flow rate varies. Overall inlet distortion, measured and described as average SAE circumferential distortion descriptor, was maintained at a value of 0.02 or less. Advantageous arrangements and orientations of the active flow control jets and the vortex generating vanes were developed using computational fluid dynamics simulations and wind tunnel experimentations.

Owens, Lewis R.↗
NASA NTRS1956

Development of flow distortions in a full-scale nacelle inlet at Mach numbers 0.63 and 1.6 to 2.0

The nature of flow-distortion development in the subsonic diffuser was determined for a typical full-scale axisymmetric nose inlet at Mach numbers to 2.0 and angles of attack to -8 degrees. Inlet design variables studied included 14 degrees and 17 degrees internalcowl lip angles conical compression surfaces with and without boundary-layer removal slots, and cone tip translation. Data presented include the inlet overall pressure recovery, mass flow, and distortion at the diffuser exit. Primary emphasis in the data, however, is placed on critical inlet operation, for which the flow distortion is traced from the inlet throat to the diffuser exit.

Piercy, Thomas G↗
NASA NTRS1973

Theoretical studies of tone noise from a fan rotor

An analytical study was made of some possible rotor alone noise sources of dipole, quadrapole and monopole characters which generate discrete tone noise. Particular emphasis is given to the tone noise caused by fan inlet flow distortion and turbulence. Analytical models are developed to allow prediction of absolute levels. Experimental data measured on a small scale fan is presented which indicates inlet turbulence interaction with a fan rotor can be a source of tone noise. Predicted and measured tone noise for the small scale rotor are shown to be in reasonable agreement.

Rao, G. V. R.↗
NASA NTRSDate not supplied

Computational Analysis of a Boundary-Layer Ingesting Tailcone Thruster Configuration Within the National Transonic Facility Using the LAVA Solver

Boundary-layer ingesting (BLI) propulsion systems are one of the many technologies currently under investigation within the aerospace community to meet NASA’s Advanced Air Transport Technology project goal of a sustainable future in aviation. To that end, an experimental campaign was conducted in the National Transonic Facility (NTF) to study the propulsion-airframe integration effects of the Boundary-Layer Ingesting Tailcone System installed in the aft portion of a 2.7% scale design of the NASA Common Research Model (CRM). This test was accompanied by a numerical simulation effort using the Launch, Ascent, and Vehicle Aerodynamics (LAVA) solver, to validate the applicability of current best-practice Reynolds-averaged Navier Stokes (RANS) models in accurately predicting the relevant flow-physics in free-air. A total of 205 steady RANS simulations were performed using LAVA, covering the range of conditions present in the NTF test matrix. Flow conditions ranged from 5- to 15-million Reynolds number, Mach numbers of 0.75, 0.80 and 0.85, and angles-of-attack between –3° and 4°. Four different mass-flow plugs were also tested to assess propulsor operating condition effects. Sensitivity to these flow conditions are analyzed in detail, especially in terms of the nacelle flow distortion which was the main subject of the study. Aerodynamic load coefficients, aftbody boundary-layer measurements and fuselage pressure tap results are also discussed, providing a wide range of validation results from this experimental campaign. This validation effort shows that current best-practices using RANS models within the LAVA solver flow solver are well suited for predicting the inlet flow distortion characteristics of novel aircraft configurations employing BLI at the aft end of the fuselage.

AATT↗
NASA NTRSDate not supplied

Computational Analysis of a Boundary Layer Ingesting Tailcone Thruster Configuration Using the LAVA Curvilinear Solver

Boundary layer ingesting (BLI) propulsion systems are one of the many technologies currently under investigation within the aerospace community to meet NASA's Advanced Air Transport Technology project goal of a sustainable future in aviation. To that end, an experimental campaign was conducted in the National Transonic Facility (NTF) wind tunnel to study the propulsion-airframe integration effects of the Boundary Layer Ingesting Tailcone System installed in the aft portion of a 2.7% scale design of the NASA Common Research Model (CRM). This test was accompanied by a numerical simulation effort using the Launch, Ascent, and Vehicle Aerodynamics (LAVA) curvilinear solver to validate the applicability of current best-practice Reynolds-Averaged Navier-Stokes (RANS) models in accurately predicting the relevant flow physics in free-air. A total of 205 steady RANS simulations were performed using LAVA, covering the range of conditions present in the NTF test matrix. Flow conditions ranged from 5- to 15-million Reynolds number, Mach numbers of 0.75, 0.80 and 0.85, and angles of attack between -3 and 4 degrees. Four different mass-flow plugs were also tested to assess propulsor operating condition effects. Sensitivity to these flow conditions is analyzed in detail, especially in terms of the nacelle flow distortion which was the main subject of the study. Aerodynamic load coefficients, aft body boundary layer measurements and fuselage pressure tap results are also discussed, providing a wide range of validation results from this experimental campaign. This validation effort shows that current best-practices using RANS models within the LAVA curvilinear solver flow solver are well suited for predicting the inlet flow distortion characteristics of novel aircraft configurations employing BLI at the aft end of the fuselage. This talk will cover the combined experimental and numerical efforts that resulted from this BLI-focused investigation.

ARMD↗
NASA NTRS2015

Investigation of Unsteady Flow Interaction Between an Ultra-Compact Inlet and a Transonic Fan

In the present study, unsteady flow interaction between an ultra-compact inlet and a transonic fan stage is investigated. Future combat aircraft require ultra-compact inlet ducts as part of an integrated, advanced propulsion system to improve air vehicle capability and effectiveness to meet future mission needs. The main purpose of the study is to advance the current understanding of the flow interaction between two different ultra-compact inlets and a transonic fan for future design applications. Both URANS and LES approaches are used to calculate the unsteady flow field and are compared with the available measured data. The present study indicates that stall inception is mildly affected by the distortion pattern generated by the inlet with the current test set-up. The numerical study indicates that the inlet distortion pattern decays significantly before it reaches the fan face for the current configuration. Numerical results with a shorter distance between the inlet and fan show that counter-rotating vortices near the rotor tip due to the serpentine diffuser affects fan characteristics significantly.

Inlet↗
NASA NTRS1972

Effects of vane-induced rotation on diffuser flow distortion in an axisymmetric mixed-compression inlet

An investigation of vane-induced flow rotation to modify distorted steady-state total-pressure patterns in the subsonic diffuser of a supersonic mixed-compression inlet was conducted. Radial static-pressure gradients generated by the rotation was the mechanism used to modify the total-pressure distributions. Significant redistribution of circumferential distortion patterns into more compatible radial patterns was realized, but flow problems near the duct walls reduced the general effectiveness of the technique. Total-pressure losses associated with the swirl vanes were slight. Limited turbulence data indicated that vane istallation resulted in reduced turbulence levels.

Meleason, E. T.↗
NASA NTRS1971

Experimental quiet engine program aerodynamic performance of fan A

The aerodynamic component test results are presented of fan A, one of two high-bypass-ratio, 1160 feet per second single-stage fans, which was designed and tested as part of the NASA Experimental Quiet Engine Program. This fan was designed to deliver a bypass pressure ratio of 1.50 with an adiabatic efficiency of 86.5% at a total fan flow of 950 lb/sec. It was tested with and without inlet flow distortion. A bypass total-pressure ratio of 1.52 and an adiabatic efficiency of 88.3% at a total fan flow of 962 lb/sec were actually achieved. An operating margin of 12.4% was demonstrated at design speed.

Giffin, R. G.↗
NASA NTRS1972

Factors affecting altitude relight performance of a double-annular ram-induction combustor

A test program was conducted to evaluate the altitude relight capabilities of a short-length, double-annular, ram-induction combustor which was designed for Mach 3 cruise operation. The use of distorted inlet-air flow profiles was tried to evaluate their effect on the relight performance. No significant improvement in altitude relight performance was obtained with this approach. A study was also made to determine the effects of the reference Mach number, the fuel temperature, and the fuel volatility (ASTM-A1 against JP-4) on the altitude relight performance. Decreasing the reference Mach number, increasing the fuel temperature, and using more volatile fuel all decrease the combustor pressure necessary for relight.

Schultz, D. F.↗
NASA NTRS1974

High-tip-speed, low-loading transonic fan stage. Part 3: Final report

Tests were conducted on a high-tip-speed, low-loading transonic fan stage to determine the performance and inlet flow distortion tolerance of the design. The fan was designed for high efficiency at a moderate pressure ratio by designing the hub section to operate at minimum loss when the tip operates with an oblique shock. The design objective was an efficiency of 86 percent at a pressure ratio of 1.5, a specific flow (flow per unit annulus area) of 42 lb/sec-sq. ft (205.1 kgm/sec-m sq), and a tip speed of 1600 ft/sec (488.6 m/sec). During testing, a peak efficiency of 84 percent was achieved at design speed and design specific flow. At the design speed and pressure ratio, the flow was 4 percent greater than design, efficiency was 81 percent, and a stall margin of 24 percent was obtained. The stall line was improved with hub radial distortion but was reduced when the stage was tested with tip radial and circumferential flow distortions. Blade-to-blade values of static pressures were measured over the rotor blade tips.

Ware, T. C.↗
NASA NTRS1975

An observation on tone cut-off in static test data from jet engine fans

The far-field acoustic data for eight full-scale fans tested at a NASA-Lewis outdoor facility were analyzed for the effect of a cut-off design feature where the number of rotor and stator blades are proportioned so that the blade-passage tone will not propagate. Inlet flow distortions, prevalent during static testing, interact with the rotor to generate tones that generally mask this cut-off phenomenon. Distortion tone properties previously evaluated are used to analyze the tones propagating from the aft fan duct. The analyzed data clearly differentiated the fans that are and are not cut-off.

Heidmann, M. F.↗
NASA NTRS1975

Propulsion system structural integration and engine integrity; Proceedings of the Symposium, U.S. Naval Postgraduate School, Monterey, Calif., September 3-6, 1974

Current problems involving the structural integrity of propulsion systems, and proposed approaches to solving them, are reviewed. Areas investigated include the stall hammershock at the engine inlet, distortion-induced vibration in fan and compressor blading, modeling engine static structures with conical-shell finite elements, design and development of low-cost, self-contained bearing lubrication systems for turbine engines, and roller bearing slip and skidding damage. Individual items are announced in this issue.

Walker, W. J.↗
NASA NTRS1976

Modal structure inferred from static far-field noise directivity

Turbofan noise directivity calculated for two directivity models was compared with experimental, blade passing frequency data from two fans at 60 and 90 percent speeds. Experimental data indicated similar directivity patterns which were well represented by a single average data curve. Calculated points using the equal amplitude model showed over-prediction near the fan axis and near the 90 degree position. Calculated points using the equal power model showed a very good match with the average data lending support to theory of equipartition of modal power from a random source such as the interaction of the rotor with inlet flow distortion. The equal modal power model also gave good agreement with individual data points.

Saule, A. V.↗
NASA NTRS1976

Modal structure inferred from static far-field noise directivity

Turbofan noise directivity calculated for two directivity models (equal modal amplitude and equal modal power or energy) was compared with experimental blade passing frequency data from two fans at 60 and 90 percent speeds. Experimental data indicated similar directivity patterns which were well represented by a single average data curve. Calculated points using the equal amplitude model showed over-prediction near the fan axis and near the 90-deg position. Calculated points using the equal power model showed a very good match with the average data lending support to theory of equipartition of modal power from a random source such as the interaction of the rotor with inlet flow distortion. The equal modal power model also gave good agreement with individual data points.

Saule, A. V.↗
NASA NTRS1980

Distribution analysis for F100(3) engine

The F100(3) compression system response to inlet circumferential distortion was investigated using an analytical compressor flow model. Compression system response to several types of distortion, including pressure, temperature, and combined pressure/temperature distortions, was investigated. The predicted response trends were used in planning future F100(3) distortion tests. Results show that compression system response to combined temperature and pressure distortions depends upon the relative orientation, as well as the individual amplitudes and circumferential extents of the distortions. Also the usefulness of the analytical predictions in planning engine distortion tests is indicated.

Walter, W. A.↗