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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 37 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRS1972

Subsonic and supersonic jet flow and acoustic characteristics and supersonic suppressors

To study the similarities and differences between subsonic and supersonic jets, velocity and impact pressure fluctuations were determined along the axis over a jet Mach number range of 0.6 to 1.4 for a 2 in. diameter convergent nozzle and for a one inch diameter jet flow. Static pressure distribution fluctuations due to shear and turbulence in the jet flow for subsonic and supersonic jets were related to acoustic radiation to the far field. Also determined were flow and acoustic characteristics of a single shroud, and multiple shroud tube and shroud suppressors for supersonic and subsonic exhaust velocities. A compressor consisting of 191 tubes and 191 shrouds decreased the primary Mach number drastically for both jet Mach numbers of 1.4 and 0.7; rms impact and static pressure fluctuations on the axis were also reduced from values existing for an equivalent area single nozzle.

Nagamatsu, H. T.↗
NASA NTRS1972

Subsonic and supersonic jets and supersonic suppressor characteristics

Convergent and parallel flow nozzles were used with room temperature air to investigate the flow and acoustic characteristics over a Mach number range of 0.6 to 1.5. For subsonic jets the maximum sound pressure levels occurred at 19.1 from the axis and the sound pressure level decreased monotonically with increasing angle. But for supersonic jets the sound pressure was nearly constant over most of angular positions. The power spectra for subsonic jets were similar with the peak occurring at 4 KHz and for a Mach 1.4 jet the peak occurred at 5 KHz. Near field sound pressure level distributions were quite similar for subsonic jets, but for supersonic jets the distributions were quite different. Mean and fluctuating velocity contours were determined for a convergent nozzle with the impact and piezoelectric impact pressure gage probes for Mach numbers of 0.6 to 1.4. A supersonic suppressor of 191 tubes and shrouds was investigated at Mach numbers of 1.4 and 0.7.

Nagamatsu, H. T.↗
NASA NTRS1976

Advanced Supersonic Technology Study: Engine Program Summary. Supersonic Propulsion: 1971 to 1976

Sustained supersonic cruise propulsion systems for military applications are studied. The J79-5 in the Mach 2 B-58; YJ93 in the Mach 3.0 B-70 and the current F101 in the B-1, are all examples of military propulsion systems and airplanes operated at sustained supersonic cruise speeds. The Mach 2.7 B2707 transport powered by GE4 turbojet engines was the only non-military, sustained supersonic cruise vehicle intended for commercial passenger service.

Krebs, J. N.↗
NASA NTRS2015

On the Comparison of the Long Penetration Mode (LPM) Supersonic Counterflowing Jet to the Supersonic Screech Jet

The authors provide a brief overview of the classic tonal screech noise problem created by underexpanded supersonic jets, briefly describing the fluid dynamic-acoustics feedback mechanism that has been long established as the basis for this well-known aeroacoustics problem. This is followed by a description of the Long Penetration Mode (LPM) supersonic underexpanded counterflowing jet phenomenon which has been demonstrated in several wind tunnel tests and modeled in several computational fluid dynamics (CFD) simulations. The authors provide evidence from test and CFD analysis of LPM that indicates that acoustics feedback and fluid interaction seen in LPM are analogous to the aeroacoustics interactions seen in screech jets. Finally, the authors propose applying certain methodologies to LPM which have been developed and successfully demonstrated in the study of screech jets and mechanically induced excitation in fluid oscillators for decades. The authors conclude that the large body of work done on jet screech, other aeroacoustic phenomena, and fluid oscillators can have direct application to the study and applications of LPM counterflowing supersonic cold flow jets.

Farr, Rebecca A.↗
NASA NTRS2015

On the Comparison of the Long Penetration Mode (LPM) Supersonic Counterflowing Jet to the Supersonic Screech Jet

Classic tonal screech noise created by under-expanded supersonic jets; Long Penetration Mode (LPM) supersonic phenomenon -Under-expanded counter-flowing jet in supersonic free stream -Demonstrated in several wind tunnel tests -Modeled in several computational fluid dynamics (CFD) simulations; Discussion of LPM acoustics feedback and fluid interactions -Analogous to the aero-acoustics interactions seen in screech jets; Lessons Learned: Applying certain methodologies to LPM -Developed and successfully demonstrated in the study of screech jets -Discussion of mechanically induced excitation in fluid oscillators in general; Conclusions -Large body of work done on jet screech, other aero-acoustic phenomenacan have direct application to the study and applications of LPM cold flow jets

Farr, Rebecca A.↗
NASA NTRSDate not supplied

NASA Supersonics Research — Leading Toward Quiet Supersonic Flight!!

- Brief history of supersonic flight - Shock waves and sonic booms - Sonic boom pressure signatures and loudness - Tools for studying quiet supersonic flight - CFD (Computational Fluid Dynamics) - Wind tunnels - Flight test - Quesst Mission - X-59 airplane - Mission profile - Flight simulations - Wind tunnel data validations - Video: Supersonic flight research leading to Quesst

sonic boom↗
NASA NTRSDate not supplied

Nasa Supersonics Research and the X-59 —To Enable Future Supersonic Travel!

- Brief history of supersonic flight - Shock waves and sonic booms - Sonic boom pressure signatures and loudness - Tools for studying quiet supersonic flight - CFD (Computational Fluid Dynamics) - Wind tunnels - Flight test - Quesst Mission - X-59 airplane - Mission profile - Flight simulations - Wind tunnel data validations - Video: Supersonic flight research leading to Quesst

sonic boom↗
NASA NTRS1973

Subsonic and supersonic jets and supersonic suppressor characteristics.

The flow and acoustic characteristics over a Mach number range of 0.6 to 1.5 were determined for convergent and parallel flow nozzles with room temperature air. Mean velocity and fluctuating impact and static pressure distributions were investigated. Near- and far-field acoustic measurements were obtained, and overall sound power levels were compared with subsonic theory of Lighthill and supersonic theory of Nagamatsu and Horvay. Flow and acoustic characteristics of single and multiple shroud supersonic suppressors were studied. With 191 tubes and shrouds the overall sound power level for a Mach 1.4 jet was reduced 15.3 and 20.5 db respectively.

Nagamatsu, H. T.↗
NASA NTRS1977

Subsonic and supersonic aerodynamic characteristics of a supersonic cruise fighter model with a twisted and cambered wing with 74 deg sweep

A wind tunnel investigation has been conducted to determine the longitudinal and lateral aerodynamic characteristics of a model of a supersonic cruise fighter configuration with a design Mach number of 2.60. The configuration is characterized by a highly swept arrow wing twisted and cambered to minimize supersonic drag due to lift, twin wing mounted vertical tails, and an aft mounted integral underslung duel-engine pod. The investigation also included tests of the configuration with larger outboard vertical tails and with small nose strakes.

Morris, O. A.↗
NASA NTRS1952

Pressure distribution and aerodynamic coefficients associated with heat addition to supersonic air stream adjacent to two-dimensional supersonic wing

The modifications in the pressure distributions and the aerodynamic coefficients associated with additions of heat to the two-dimensional supersonic in viscid flow field adjacetnt to the lower surface of of a 5-percent-thickness symmetrical circular-arc wing are presented in this report. The pressure distributions are obtained by the use of graphical method which gives the two-dimensional supersonic inviscid flow field obtained with moderate heat addition. The variation is given of the lift-drag ratio and of the aerodynamic coefficients of lift, drag, and moment with free stream Mach number, angle of attack, and parameters defining extent and amount of heat addition. The six graphical solutions used in this study included Mach numbers of 3.0 and 5.0 and angles of attack of 0 degrees and 2 degrees.

Pinkel, I Irving↗
NASA NTRS1950

Theoretical lift and damping in roll at supersonic speeds of thin sweptback tapered wings with streamwise tips, subsonic leading edges, and supersonic trailing edges

On the basis of linearized supersonic-flow theory, generalized equations were derived and calculations made for the lift and damping in roll of a limited series of thin sweptback tapered wings. Results are applicable to wings with streamwise tips and for a range of supersonic speeds for which the wing is wholly contained between the Mach cones springing from the wing apex and from the trailing edge of the root section. A further limitation is that the tip Mach lines may not intersect on the wing. For the portion of the wing external to the Mach cones springing from the leading edge of the wing tips, the pressure distributions for lift and roll previously obtained for the triangular wing are valid. For the portion of the wing contained within the wing-tip Mach cones a satisfactory approximation to the exact pressure distribution was obtained by application of a point-source-distribution method developed in NACA-TN-1382. A series of design curves are presented which permit rapid estimation of the lift-curve slope and damping-in-roll derivative for given values of aspect ratio, taper ratio, Mach number, and leading-edge sweep. (author)

Malvestuto, Frank S , Jr↗
NASA NTRS1952

Airfoil profiles for minimum pressure drag at supersonic velocities -- general analysis with application to linearized supersonic flow

A theoretical investigation is made of the airfoil profile for minimum pressure drag at zero lift in supersonic flow. In the first part of the report a general method is developed for calculating the profile having the least pressure drag for a given auxiliary condition, such as a given structural requirement or a given thickness ratio. The various structural requirements considered include bending strength, bending stiffness, torsional strength, and torsional stiffness. No assumption is made regarding the trailing-edge thickness; the optimum value is determined in the calculations as a function of the base pressure. To illustrate the general method, the optimum airfoil, defined as the airfoil having minimum pressure drag for a given auxiliary condition, is calculated in a second part of the report using the equations of linearized supersonic flow.

Chapman, Dean R↗
NASA NTRS1997

Future Directions of Supersonic Combustion Research: Air Force/NASA Workshop on Supersonic Combustion

The Air Force Office of Scientific Research, the Air Force Wright Laboratory Aero Propulsion and Power Directorate, and the NASA Langley Research Center held a joint supersonic combustion workshop on 14-16 May 1996. The intent of this meeting was to: (1) examine the current state-of-the-art in hydrocarbon and/or hydrogen fueled scramjet research; (2) define the future direction and needs of basic research in support of scramjet technology; and (3) when appropriate, help transition basic research findings to solve the needs of developmental engineering programs in the area of supersonic combustion and fuels. A series of topical sessions were planned. Opening presentations were designed to focus and encourage group discussion and scientific exchange. The last half-day of the workshop was set aside for group discussion of the issues that were raised during the meeting for defining future research opportunities and directions. The following text attempts to summarize the discussions that took place at the workshop.

Tishkoff, Julian M.↗
NASA NTRS2012

Supersonic Retropropulsion Experimental Results from the NASA Ames 9- x 7-Foot Supersonic Wind Tunnel

Supersonic retropropulsion was experimentally examined in the Ames Research Center 9x7-Foot Supersonic Wind Tunnel at Mach 1.8 and 2.4. The experimental model, previously designed for and tested in the Langley Research Center Unitary Plan Wind Tunnel at Mach 2.4, 3.5 and 4.6, was a 5-in diameter 70-deg sphere-cone forebody with a 9.55-in long cylindrical aftbody. The forebody was designed to accommodate up to four 4:1 area ratio nozzles, one on the model centerline and the other three on the half radius spaced 120-deg apart. Surface pressure and flow visualization were the primary measurements, including high-speed data to investigate the dynamics of the interactions between the bow and nozzle shocks. Three blowing configurations were tested with thrust coefficients up to 10 and angles of attack up to 20-deg. Preliminary results and observations from the test are provided

Berry, Scott A.↗
NASA NTRS1971

Mixing of supersonic jets including the effects of supersonic jets including the effects of transverse pressure gradient using difference methods

The usual boundary-layer equations describing the steady-state mixing of parallel jets are supplemented by the momentum equation in the direction normal to the flow. This allows detailed computation of the flow field in the mixing region and simultaneous computation of the outer inviscid flow. An explicit and an implicit finite difference scheme have been developed and applied in several illustrative examples. The examples include mixing of planar and axisymmetric supersonic jets of different composition with both matched and unmatched static pressures. Numerical results were compared with available experimental data obtained for the unmatched pressure case.

Kurkov, A. P.↗