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Burstadt, P. L.

Publications and source records attributed to Burstadt, P. L..

17 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1996

Experimental Investigation of the Performance of a Mach-2.7 Two-dimensional Bifurcated Duct Inlet with 30 Percent Internal Contraction

An experimental study was conducted to determine the performance of a two-dimensional, mixed-compression bifurcated duct inlet system designed for a free-stream Mach number of 2.7. Thirty percent of the supersonic area contraction occurred internally. A movable ramp was used to vary the contraction ratio for off-design operation. Boundary layer bleed regions were located on the cowl, centerbody, and sidewall surfaces. There were also provisions for vortex generators on the cowl and centerbody of the subsonic diffuser. Data were obtained over the Mach number range of 2.0 to 2.8 and at angles of yaw from 0 deg. to the maximum value prior to inlet un-start. The test at Mach 2.8 was to obtain data for an over- speed condition. The Reynolds number varied from 2.5 to 2.3 million/ft for Mach numbers above 2.5. At Mach numbers of 2.5 and lower, the Reynolds number was set at 2.5 million/ft. Bleed patterns, vortex generator patterns, and ramp position were varied, and three inlet configurations were selected for more extensive study. Two of these configurations had self-starting capability. The self-starting configuration that was developed produced 89 percent total pressure recovery at the compressor face station with 6.8 percent total bleed. The compressor face distortion was about 16 percent. Vortex generators were extremely effective in re-distributing flow but were not as effective in reducing distortion. Excellent flow symmetry was achieved between the separated halves of the inlet, and twin-duct instability was not observed. The ramp tip shock was steeper than expected. This caused the cowl lip shock to be reflected from the ramp instead of being cancelled at the shoulder. However, peak recovery at the throat was still obtained with the ramp near the design position.

Wasserbauer, J. F.↗
NASA NTRS1984

Flow visualization and interpretation of visualization data for deflected thrust V/STOL nozzles

Flow visualization studies were made for four deflected thrust nozzle models at subsonic speeds. Based on topological rules and the assumption that observed streaks constitute continuous vector fields, available visualization pictures are interpreted and flow patterns on interior surfaces of the nozzles are synthesized. In particular, three dimensional flow structure and separations are discussed. From the synthesized patterns, the overall features of the flow field in a given nozzle can be approximately perceived.

Kao, H. C.↗
NASA NTRS1984

Flow visualization and interpretation of visualization data for deflected thrust V/STOL nozzles

Flow visualization studies were made for four deflected thrust nozzle models at subsonic speeds. Based on topological rules and the assumption that observed streaks constitute continuous vector fields, available visualization pictures are interpreted and flow patterns on interior surfaces of the nozzles are synthesized. In particular, three dimensional flow structure and separations are discussed. From the synthesized patterns, the overall features of the flow field in a given nozzle can be approximately perceived. Previously announced in STAR as N84-14147

Kao, H. C.↗
NASA NTRS1983

Experimental results of a deflected thrust V/STOL nozzle research program

Four deflected thrust nozzle concepts, designed to operate at the low pressure ratio typical of high bypass-ratio turbofan engines for medium speed (subsonic) V/STOL aircraft, were studied. Maps of overall performance characteristics and exit velocity distributions are used to highlight similarities and differences between the four concepts. Analytically determined secondary flows at the exit of a 90 deg circular pipe bend are compared with the experimental results from the more complex three dimensional geometries. The relative impact of total-pressure losses and secondary flows on nozzle thrust coefficient is addressed by numerical integration of exit velocity measurements.

Burstadt, P. L.↗
NASA NTRS1983

Experimental results of a deflected thrust V/STOL nozzle research program

Four deflected thrust nozzle concepts, designed to operate at the low pressure ratio typical of high bypass-ratio turbofan engines for medium speed (subsonic) V/STOL aircraft, were studied. Maps of overall performance characteristics and exit velocity distributions are used to highlight similarities and differences between the four concepts. Analytically determined secondary flows at the exit of a 90 deg circular pipe bend are compared with the experimental results from the more complex three dimensional geometries. The relative impact of total-pressure losses and secondary flows on nozzle thrust coefficient is addressed by numerical integration of exit velocity measurements. Previously announced in STAR as N83-25657

Burstadt, P. L.↗
NASA NTRS1976

A method to account for variation of average compressor inlet pressure during instantaneous distortion analyses

Instantaneous distortion analyses compare a time-varying value of an index (or 'surge margin used up') with a critical level (or 'available surge margin' of the compressor) to determine inlet-engine compatibility. Unless freestream conditions or propulsion system controls are changing, it is generally assumed that the available surge margin of the compressor is accurately determined from the steady-state operating point. Results are presented which show that variations of average compressor inlet pressure may occur without changes in freestream conditions or propulsion system controls. The volume dynamics of the compressor will cause these pressure variations to be attenuated and delayed by the time they reach the exit. This will cause the compressor pressure ratio (and available surge margin) to vary with time. A method is presented to calculate the available surge margin as a function of time and incorporate it into an instantaneous distortion analysis. Results show that inlet pressure variations which cause only a small change at the compressor exit can cause a significant variation in the available surge margin.

Burstadt, P. L.↗
NASA NTRS1975

Aerodynamic characteristics of MSFC model 454 of the 142 inch solid rocket booster tested in the LeRC 10 foot SWT at Mach numbers of 2.0 and 2.7 (SA6F)

A 2.112 percent scale Space Shuttle Solid Rocket Booster (SRB) was tested in a ten foot, supersonic wind tunnel. The test Mach numbers were 2.0 and 2.7. Test angles of attack were from minus 5 degrees to plus 185 degrees. The Reynolds numbers ranged from 0.514 to 2.81 million per foot. Test roll angles were 0, 22.5, 45, 90, and 135 degrees. The following configurations were tested: (1) SRB without external protuberances, (2) SRB with an electrical tunnel and a thrust attachment structure, (3) SRB with two engine shroud strakes, (4) SRB with eight engine shroud strakes, and (5) SRB with an electrical tunnel, thrust attachment structure, eight engine shroud strakes, and separation motors.

Johnson, J. D.↗
NASA NTRS1975

Propulsion systems technology

The selection of components, control design, simulation, and tests are examined for designing supersonic propulsion systems. Inlet-engine combination under electronic control are shown. Design and test results from the integrated control for afterburning turbofan and supersonic cruise inlet are presented.

Hiller, K. W.↗
NASA NTRS1974

Instantaneous distortion in a Mach 2.5, 40-percent-internal-contraction inlet and its effect on turbojet stall margin

An experimental wind tunnel investigation was conducted to determine the effects of time-variant distortions produced in a supersonic inlet on a J85-GE-13 turbojet engine. Results are presented principally in terms of instantaneous distortion amplitudes and total-pressure contours measured through compressor stall. They indicate that, although a time-averaged distortion may be far from a stall-inducing value, corresponding instantaneous distortion amplitudes can approach or exceed this value. Seven engine stall events were studied. In six of these events instantaneous distortion peaks of sufficient magnitude to cause stall were observed in the time period shortly before stall was detected.

Burstadt, P. L.↗
NASA NTRS1974

Effect of screen-induced total-pressure distortion on axial-flow compressor stability

An experimental investigation was made to determine the effects of screen-induced total-pressure distortions on two J85-GE-13 turbojet engines. Results were compared to those from a previous program run with a third engine. All compressors were found to be sensitive to a critical angle of circumferential distortion equal to 60 deg., and they all adhered closely to the parallel compressor model. The sensitivity of compressor exit pressure to virtually any type of distortion pattern can be determined by defining stall lines for undistorted, hub radial distorted, and tip radial distorted inflows. The effect of multiple sectors of circumferential distortion is defined.

Calogeras, J. E.↗
NASA NTRS1974

Formulation of a distortion index based on peak compressor pressure ratios

In order to effectively use a compressor face total-pressure distortion index as a measure of inlet-engine compatibility, a correlation of distortion amplitude with stall margin must be developed with minimal scatter. A recent analysis of data recorded in extensive distortion screen tests with the J85-GE-13 turbojet engine has resulted in a correlation based on compressor discharge pressure ratioed to the minimum pressure at the compressor face. Simply by determining compressor stall lines with a single hub radial distortion pattern, a single tip radial pattern, and with undistorted inflow, the overall compressor pressure ratio at stall for even the most complex distortion pattern was found to be predictable.

Calogeras, J. E.↗
NASA NTRS1974

Formulation of a distortion index based on peak compressor pressure ratios

In order to effectively use a compressor face total-pressure distortion index as a measure of inlet-engine compatibility, a correlation of distortion amplitude with stall margin must be developed with minimal scatter. A recent analysis of data recorded in extensive distortion screen tests with the J85-GE-13 turbojet engine has resulted in a correlation based on compressor discharge pressure ratioed to the minimum pressure at the compressor face. Simply by determining compressor stall lines with a single hub radial distortion pattern, a single tip radial pattern, and with undistorted inflow, the overall compressor pressure ratio at stall for even the most complex distortion pattern was found to be predictable.

Calogeras, J. E.↗