Search NASA⌕ Search

SEARCH · Search NASA

Results for “SHOCK ATTACHMENT”

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.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 55 records · Page 3
All sourcesNASA NTRSDOE OSTI
NASA NTRS1998

Experimental and Theoretical Study of Flow Fields Around Ducted-Nacelle Models

The flow field near four small-scale ducted-nacelle bodies of revolution has been analytically and experimentally studied to determine exterior and interior mass-flow characteristics, and to measure flow-field overpressures generated by the nacelle's forebody shape. Four nacelle models with the same profile, but of different sizes, were used in the study. Shadowgraph pictures showed inlet shocks attached to the cowl lip (indicating unchoked flow) on all four models, at all the test Mach numbers, through an angle of attack range of 0.0 to 6.0 degrees. Pressure signatures measured in the flow field of the largest of the four nacelle models were compared with those predicted by corrected and uncorrected Whitham theory. At separation distances greater than 3.0 to 4.0 inlet diameters, good agreement was found. Poorer agreement was found at extreme near-field separation distances, but this was attributed to pressure-gage limitations and probe-flow field interactions. The overall favorable results supported a conclusion that corrected Whitham theory was sufficiently accurate to make the nacelle-wing interference-lift code useful for sonic-boom analysis and the preliminary design of supersonic-cruise conceptual aircraft.

Mack, Robert J.↗
NASA NTRS1961

Measurement of Flow Angularity at Supersonic and Hypersonic Speeds with the Use of a Conical Probe

The characteristics of a 40 deg half-angle cone for measuring flow angularity were determined experimentally. Tests were conducted at a Mach number of 21 in helium with check points at Mach number 3.55 in air for angles of pitch up to 5 deg. The air tests confirmed theoretical indications of small or negligible Mach number and test-medium effects for the case of air and helium. The instrument is capable of measuring flow angularity at high Reynolds numbers and speeds greater than that necessary for shock attachment to within (+/-)1/3 deg.

Swalley, Frank E.↗
NASA NTRS1995

Aerodynamic Performance and Flow-Field Characteristics of Two Waverider-Derived Hypersonic Cruise Configurations

The component integration of a class of hypersonic high-lift configurations known as waveriders into hypersonic cruise vehicles was evaluated. A wind-tunnel model was developed which integrates realistic vehicle components with two waverider shapes, referred to as the straight-wing and cranked-wing shapes. Both shapes were conical-flow-derived waveriders for a design Mach number of 4.0. Experimental data and limited computational fluid dynamics (CFD) predictions were obtained over a Mach number range of 1.6 to 4.63 at a Reynolds number of 2.0 x 10(exp 6) per foot. The CFD predictions and flow visualization data confirmed the shock attachment characteristics of the baseline waverider shapes and illustrated the waverider flow-field properties. Experimental data showed that no significant performance degradations, in terms of maximum lift-to-drag ratios, occur at off-design Mach numbers for the waverider shapes and the integrated configurations. A comparison of the fully-integrated waverider vehicles to the baseline shapes showed that the performance was significantly degraded when all of the components were added to the waveriders, with the most significant degradation resulting from aftbody closure and the addition of control surfaces. Both fully-integrated configurations were longitudinally unstable over the Mach number range studied with the selected center of gravity location and for unpowered conditions. The cranked-wing configuration provided better lateral-directional stability characteristics than the straight-wing configuration.

Cockrell, Charles E., Jr.↗
NASA NTRS1993

Interpretation of Waverider Performance Data Using Computational Fluid Dynamics

A computational study was conducted to better understand experimental results obtained from wind tunnel tests of a Mach 4 waverider model and a comparative reference configuration. The experimental results showed that the performance of the reference configuration was slightly better than that of the waverider model. These results contradict waverider design theory, which suggests that a waverider optimized for maximum lift-to-drag should provide better performance than any other non-waverider configuration at a given design point, especially at hypersonic speeds. The computational results showed that the predicted surface pressure values and the integrated lift and drag coefficients from the pressure distributions were much lower for the reference model than for the flat-top model, due to the reference model bottom surface having a slight expansion. The lift-to-drag ratios for the flat-top model were higher due to a relatively low drag for the same amount of lift. These results indicate that the performance advantage of the reference model was due to the shape of the bottom surface and not due to the flat top surface. The results also showed that the reference model exhibited the same shock attachment characteristics as the waverider because the planform shapes were identical. CFD predictions show that the planform shape gives the waverider an advantage in performance over conventional hypersonic vehicles and that altering the bottom surface of a waverider does not cause significant performance degradation.

Cockrell, Charles E., Jr.↗
NASA NTRS2004

Estimating Basic Preliminary Design Performances of Aerospace Vehicles

Aerodynamics and Performance Estimation Toolset is a collection of four software programs for rapidly estimating the preliminary design performance of aerospace vehicles represented by doing simplified calculations based on ballistic trajectories, the ideal rocket equation, and supersonic wedges through standard atmosphere. The program consists of a set of Microsoft Excel worksheet subprograms. The input and output data are presented in a user-friendly format, and calculations are performed rapidly enough that the user can iterate among different trajectories and/or shapes to perform "what-if" studies. Estimates that can be computed by these programs include: 1. Ballistic trajectories as a function of departure angles, initial velocities, initial positions, and target altitudes; assuming point masses and no atmosphere. The program plots the trajectory in two-dimensions and outputs the position, pitch, and velocity along the trajectory. 2. The "Rocket Equation" program calculates and plots the trade space for a vehicle s propellant mass fraction over a range of specific impulse and mission velocity values, propellant mass fractions as functions of specific impulses and velocities. 3. "Standard Atmosphere" will estimate the temperature, speed of sound, pressure, and air density as a function of altitude in a standard atmosphere, properties of a standard atmosphere as functions of altitude. 4. "Supersonic Wedges" will calculate the free-stream, normal-shock, oblique-shock, and isentropic flow properties for a wedge-shaped body flying supersonically through a standard atmosphere. It will also calculate the maximum angle for which a shock remains attached, and the minimum Mach number for which a shock becomes attached, all as functions of the wedge angle, altitude, and Mach number.

Luz, Paul L.↗
NASA NTRS1972

A method for calculating aerodynamic heating on sounding rocket tangent ogive noses

A method is presented for calculating the aerodynamic heating and shear stresses at the wall for tangent ogive noses that are slender enough to maintain an attached nose shock through that portion of flight during which heat transfer from the boundary layer to the wall is significant. The lower entropy of the attached nose shock combined with the inclusion of the streamwise pressure gradient yields a reasonable estimate of the actual flow conditions. Both laminar and turbulent boundary layers are examined and an approximation of the effects of (up to) moderate angles-of-attack is included in the analysis. The analytical method has been programmed in FORTRAN 4 for an IBM 360/91 computer.

Wing, L. D.↗
NASA NTRS1973

A method for calculating aerodynamic heating on sounding rocket tangent ogive noses.

A method is presented for calculating the aerodynamic heating and shear stresses at the wall for tangent ogive noses that are slender enough to maintain an attached nose shock through that portion of flight during which heat transfer from the boundary layer to the wall is significant. The lower entropy of the attached nose shock combined with the inclusion of the streamwise pressure gradient yields a reasonable estimate of the actual flow conditions. Both laminar and turbulent boundary layers are examined and an approximation of the effects of (up to) moderate angles-of-attack is included in the analysis. The analytical method has been programmed in FORTRAN IV for an IBM 360/91 computer.

Wing, L. D.↗
NASA NTRS1992

Oblique shock wave with sweep

An attached planar, oblique shock with sweep is investigated for the inviscid flow of a perfect gas. The ratio of specific heats, freestream Mach number, and wedge angle in the plane of the freestream velocity are prescribed, with the sweep angle as a free parameter. Explicit relations are provided for jump and detachment conditions. A number of trends, some nonintuitive, are discussed, e.g., the downstream Mach number may increase with sweep.

Emanuel, G.↗
NASA NTRS2010

The Prediction of Broadband Shock-Associated Noise from Dualstream and Rectangular Jets Using RANS CFD

Supersonic jets operating off-design produce broadband shock-associated noise. Broadband shock-associated noise is characterized by multiple broadband peaks in the far-field and is often the dominant source of noise towards the sideline and upstream direction relative to the jet axis. It is due to large scale coherent turbulence structures in the jet shear layers interacting with the shock cell structure. A broadband shock-associated noise model recently developed by the authors predicts this noise component from solutions to the Reynolds averaged Navier-Stokes equations using a two-equation turbulence model. The broadband shock-associated noise model is applied to dualstream and rectangular nozzles operating supersonically, heated, and off-design. The dualstream jet broadband shock-associated noise predictions are conducted for cases when the core jet is supersonic and the fan jet is subsonic, the core jet is subsonic and the fan jet is supersonic, and when both jet streams operate supersonically. Rectangular jet predictions are shown for a convergent-divergent nozzle operating both over- and under-expanded for cold and heated conditions. The original model implementation has been heavily modified to make accurate predictions for the dualstream jets. It is also argued that for over-expanded jets the oblique shock wave attached to the nozzle lip contributes little to broadband shock-associated noise. All predictions are compared with experiments.

Miller, Steven A.↗
NASA NTRS1980

Pyrotechnic shock at the orbiter/external tank forward attachment

During the initial certification test of the forward structural attachment of the space shuttle orbiter to the external tank, pyrotechnic shock from actuation of the separation device resulted in structural failure of the thermal protection tiles surrounding the attachment. Because of the high shock associated with the separation bolt, the development of alternative low shock separation designs was initiated. Two concepts that incorporate a 5.08 centimeter frangible nut as the release device were developed and tested.

Rogers, W. F.↗
NASA NTRS1977

Computation of the inviscid supersonic flow about cones at large angles of attack by a floating discontinuity approach

The technique of floating shock fitting is adapted to the computation of the inviscid flowfield about circular cones in a supersonic free stream at angles of attack that exceed the cone half-angle. The resulting equations are applicable over the complete range of free-stream Mach numbers, angles of attack and cone half-angles for which the bow shock is attached. A finite difference algorithm is used to obtain the solution by an unsteady relaxation approach. The bow shock, embedded cross-flow shock, and vortical singularity in the leeward symmetry plane are treated as floating discontinuities in a fixed computational mesh. Where possible, the flowfield is partitioned into windward, shoulder, and leeward regions with each region computed separately to achieve maximum computational efficiency. An alternative shock fitting technique which treats the bow shock as a computational boundary is developed and compared with the floating-fitting approach. Several surface boundary condition schemes are also analyzed.

Daywitt, J.↗
NASA NTRS1977

Supersonic flow about circular cones at large angles of attack - A floating discontinuity approach

The technique of floating shock fitting is adapted to the computation of the inviscid flowfield about circular cones in a supersonic free stream at angles of attack that exceed the cone half-angle. In those regions in which the governing conical equations are mixed elliptic-hyperbolic, the fully hyperbolic form is obtained by the addition of the temporal derivative. The resulting equations are applicable over the complete range of free-stream Mach numbers, angles of attack and cone half-angles for which the bow shock is attached. An explicit finite-difference algorithm is used to obtain the solution by an unsteady relaxation approach. The bow shock, embedded crossflow shock, and vortical singularity in the leeward symmetry plane are all treated as floating discontinuities in a fixed computational mesh. The method yields excellent results for the bow and embedded shocks, however, the solution in the leeward symmetry plane exhibits viscous-like effects and does not appear to adequately predict the behavior of the vortical singularity.

Daywitt, J.↗
NASA NTRS1974

Flight test measurements and analysis of sonic boom phenomena near the shock wave extremity

A NASA flight test program conducted during the summer and fall of 1970 was devoted to investigating sonic boom phenomena near caustics formed by steady flight near the threshold Mach number, during accelerations, and at the lateral extremes of the ground carpet. The vertical extent of the shock waves attached to near-sonic airplanes was also studied. The flights were conducted over the 1500 ft instrumented BREN tower so that vertical surveys through the shock waves were measured. These data on caustic phenomena near the shock wave extremity were analyzed in detail and compared with theoretical results. Amplifications of shock wave strength varied from 2 to 5 during longitudinal accelerations, from 1 to 1.8 during steady threshold Mach number flight, and up to 3 for small inadvertent accelerations during flight near the threshold Mach number.

Haglund, G. T.↗
NASA NTRS1990

Can shock waves on helicopter rotors generate noise? - A study of the quadrupole source

An analysis has previously established that local shock surfaces attached to helicopter rotor blades moving at high subsonic speeds are potent noise generators; in pursuit of this insight, a novel formulation is presented for the prediction of the noise of a deformable shock, whose area changes as a function of the azimuthal position of the blade. The derivation of this formulation has its basis in a mapping of the moving shock to a time-independent region. In virtue of this mapping, the implementation of the main result on a computer becomes straightforward enough for incorporation into the available rotor-noise prediction code. A problem illustrating the importance of rotor shocks in the generation of high-intensity noise is presented.

Farassat, F.↗
NASA NTRS1973

Flight test measurements and analysis of sonic boom phenomena near the shock wave extremity

The sonic boom flight test program conducted at Jackass Flats, Nevada, during the summer and fall of 1970 consisted of 121 sonic-boom-generating flights over the 1500 ft instrumented BREN tower. This test program was designed to provide information on several aspects of sonic boom, including caustics produced by longitudinal accelerations, caustics produced by steady flight near the threshold Mach number, sonic boom characteristics near lateral cutoff, and the vertical extent of shock waves attached to near-sonic airplanes. The measured test data, except for the near-sonic flight data, were analyzed in detail to determine sonic boom characteristics for these flight conditions and to determine the accuracy and the range of validity of linear sonic boom theory. The caustic phenomena observed during the threshold Mach number flights and during the transonic acceleration flights are documented and analyzed in detail. The theory of geometric acoustics is shown to be capable of predicting shock wave-ground intersections, and current methods for calculating sonic boom pressure signature away from caustics are shown to be reasonably accurate.

Haglund, G. T.↗
NASA NTRS1997

The Effect of Three-Dimensional Freestream Disturbances on the Supersonic Flow Past a Wedge

The interaction between a shock wave (attached to a wedge) and small amplitude, three-dimensional disturbances of a uniform, supersonic, freestream flow are investigated. The paper extends the two-dimensional study of Duck et al, through the use of vector potentials, which render the problem tractable by the same techniques as in the two-dimensional case, in particular by expansion of the solution by means of a Fourier-Bessel series, in appropriately chosen coordinates. Results are presented for specific classes of freestream disturbances, and the study shows conclusively that the shock is stable to all classes of disturbances (i.e. time periodic perturbations to the shock do not grow downstream), provided the flow downstream of the shock is supersonic (loosely corresponding to the weak shock solution). This is shown from our numerical results and also by asymptotic analysis of the Fourier-Bessel series, valid far downstream of the shock.

Duck, Peter W.↗