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Showing papers on "Supersonic speed published in 1968"


Journal ArticleDOI
TL;DR: In this paper, the authors describe a study concerning the sonic injection of a gaseous jet through a transverse slot nozzle in a wall into an external flow which is uniform outside of a turbulent boundary layer.
Abstract: The paper describes a study concerning the sonic injection of a gaseous jet through a transverse slot nozzle in a wall into an external flow which is uniform outside of a turbulent boundary layer. An analytic model of the flowfield has been constructed in which conservation of momentum is applied to a control volume at the jet nozzle exit. A series of flat-plate experiments was conducted with normal, sonic jets at external flow Mach numbers of 2.61, 3.50, and 4.54. Pressure data near separation and the plateau were in agreement with existing correlations. Comparisons of the trends predicted by the analysis with two-dimensional force data from these experiments and from other sources showed good agreement. Values of amplification factor, the upstream interaction force plus the jet thrust divided by the vacuum thrust of a sonic jet, of 2.9 to 3.2 were measured. The amplification factor is relatively insensitive to variations in external flow Mach number and variations in injectant gas properties. A correlation of data obtained from experiments with finite-span slots demonstrates that the effective jet penetration height and the slot span are the important characteristic dimensions of such flowfields.

238 citations


Proceedings ArticleDOI
01 Aug 1968
TL;DR: Energy state approximation for supersonic aircraft performance optimization with extension to maximum range problems was studied in this paper, where the authors compared the complexity of complex dynamic models with simple dynamic models.
Abstract: Energy state approximation for supersonic aircraft performance optimization with extension to maximum range problems, noting comparison with complex dynamic models

232 citations


Journal ArticleDOI
TL;DR: In this paper, a new method for numerical solution of time-dependent problems in several space dimensions is presented, which is applicable to low-speed (incompressible) flows, to high speed (supersonic) flows and to all flow speeds in between.

206 citations


Journal ArticleDOI
TL;DR: In this article, the aerodynamic influence coefficients have been extended to the subsonic flow regime and applied to the design of wing camber surfaces in the presence of a body.
Abstract: The method of aerodynamic influence coefficients has proved to be an effective tool for the analysis and design of wings, bodies, and wing-body combinations at supersonic speeds. This paper describes the extension of this method into the subsonic flow regime, and correlates the theory with experiment over a wide speed range. The method may be applied to the calculation of the pressures and forces acting on arbitrary wing-body combinations in steady flight, including aeroelastic effects, and to the design of wing camber surfaces in the presence of a body. Nomenclature A = aspect ratio b = span c =. chord C = aerodynamic coefficient d = distance / = singularity distribution function F = distribution function K = constant I = body length L — panel sweep M = Mach number r = radius Re = real part u, v, w = perturbation velocities U = freestream velocity x, y, z = Cartesian coordinates a. = angle of attack 0 = angular coordinate, panel inclination X = taper ratio A = leading edge sweep £, rj = integration variables

154 citations


ReportDOI
01 Apr 1968
TL;DR: In this paper, the pressure fluctuations beneath equilibrium boundary layers at subsonic and supersonic speeds are reviewed and empirical formulae are presented for the intensity, spectra, and cross spectra of the fluctuations.
Abstract: : The pressure fluctuations beneath equilibrium boundary layers at subsonic and supersonic speeds are reviewed Empirical formulae are presented for the intensity, spectra, and cross spectra of the fluctuations The formulae are intended for direct use in structural response calculations A simple theoretical model to predict intensities at supersonic speeds is put forward, and the effects of nonequilibrium boundary layers are discussed in general terms An appendix gives theoretical justification for negative exponential correlation curves at large spacings (Author)

63 citations



Journal ArticleDOI
01 Jan 1968
TL;DR: In this article, a method for rational design of lifting vehicles intended to cruise at speeds between M = 4 and 7 was proposed, and the lifting efficiencies of such shapes may be investigated and optimised.
Abstract: A method is proposed which may be used as a basis for the rational design of lifting vehicles intended to cruise at speeds between M=4 and 7. In order to construct shapes whose behaviour in an inviscid fluid stream may be calculated exactly, streamsurfaces chosen according to certain rules from axisymmetric flow fields are replaced by solid boundaries. Combinations of such streamsurfaces may be built up piecemeal into complete aircraft configurations, and it is shown that the method is sufficiently flexible to allow considerable freedom of choice in the disposition of lift and volume. Methods are indicated whereby the lifting efficiencies of such shapes may be investigated and, to some extent, optimised.

59 citations



Journal ArticleDOI
TL;DR: In this article, a study of normal and lateral spray penetration for small diameter, high pressure, liquid jets issuing at an angle to a uniform supersonic stream was carried out in a 4-in. by 4.in. blow down wind tunnel, and the results showed that the spray width behind the jet was proportional to the jet diameter with only a weak dependence on the injection pressure ratio.
Abstract: : The paper reports a study of normal and lateral spray penetration for small diameter, high pressure, liquid jets issuing at an angle to a uniform supersonic stream. The experimental program described was carried out in a 4-in. by 4-in. blow down supersonic wind tunnel. The flow field is observed by means of a schlieren system, and the spray distribution is indicated by the light scattered by the liquid droplets. The data on normal penetration, in good agreement with data inferred from other investigations, indicate that a single-parameter correlation exists between the properly nondimensionalized penetration height and the injection pressure ratio. Injecting the coolant at a forward angle to the flow produces no substantial change in the penetration height. The data on lateral penetration show the spray width behind the jet to be proportional to the jet diameter with only a weak dependence on the injection pressure ratio. Analytical models proposed by previous investigators are critically examined in light of the results. No single model leads to a proper scaling law for both normal and lateral penetration. (Author)

49 citations


Journal ArticleDOI
TL;DR: The evolution of the aerodynamic shape is shown to lead to a complete spectrum of major types of aircraft, supplemented by special types designed for particular purposes, in a first-order framework for design procedures aimed at achieving a given flight performance.

46 citations


01 May 1968
TL;DR: The FORTRAN program for approximate calculation of supersonic ideal gas flow past blunt bodies with sonic corners shows good agreement with prior work on this topic.
Abstract: FORTRAN program for approximate calculation of supersonic ideal gas flow past blunt bodies with sonic corners

Proceedings ArticleDOI
01 Oct 1968
TL;DR: Supersonic aerodynamic design tools, discussing technological application of high speed computer and limitations are discussed in this article, where the authors also discuss the limitations of high-speed computer and its application.
Abstract: Supersonic aerodynamic design tools, discussing technological application of high speed computer and limitations

Journal Article
TL;DR: In this paper, a numerical procedure for solving the problem of steady supersonic inviscid flow around smooth conical bodies is presented by solving the elliptic partial differential equations that define the conical flow between the body and the shock.
Abstract: : A numerical procedure for solving the problem of steady supersonic inviscid flow around smooth conical bodies is presented. Results are obtained by solving the elliptic partial differential equations that define the conical flow between the body and the shock. Results are given for circular cones up to moderately high relative incidences, including some cases for incidences beyond a critical value at which the entropy singularity moves from the surface. Also presented are a few results for elliptic cones at zero and non-zero incidence, as well as results for another conical body whose cross section is defined by a fourth order even cosine Fourier series. The applicability of the method can be limited by the entropy singularity moving too far away from the surface by the flow field containing regions of locally conically supersonic flow, or by the shock wave approaching very close to the Mach wave. Comparison of results shows excellent agreement with other theoretical methods and also with experimental results. The method is efficient in computer time.

Journal ArticleDOI
TL;DR: In this article, the authors studied the acoustic field of small, cold supersonic jets by shadowgraph visualization and showed the importance of choosing an optimum film position for the visualization of sound waves.
Abstract: The acoustic field of small, cold supersonic jets has been studied by shadowgraph visualization. Preliminary experiments showed the importance of choosing an optimum film position for the visualization of sound waves. The interpretation of shadowgraphs is discussed. The apparent radiation field from the supersonic region of the flow is of relatively high frequency. In this field the order of dominance of the sound sources in the present case is (i) radiation from the nozzle, (ii) radiation from shock‐turbulence interaction in the flow, and (iii) Mach‐wave radiation. When the jet flow is deflected, the jet‐deflector interaction becomes the principal noise source. The significance of the nozzle radiation is believed to result from the small scale of the jet used and demonstrates the necessity for care when extrapolating model scale acoustic data. The comparative in significance of Mach‐wave radiation is probably due to the lower exhaust velocity of the cold jets. It is suggested that Mach reflection of spherical waves may be an important factor in the generation of “Mach waves.”

Journal ArticleDOI
TL;DR: In this paper, the authors reported the results of experiments conducted in the 20in. supersonic wind tunnel at the Jet Propulsion Laboratory of the California Institute of Technology on the effect of heat transfer (cooling) on the size of single spherical roughness elements required to produce a minimum transition Reynolds number on a 43in. 10° cone.
Abstract: The paper reports the results of experiments conducted in the 20in. supersonic wind tunnel at the Jet Propulsion Laboratory of the California Institute of Technology on the effect of heat transfer (cooling) on the size of single spherical roughness elements required to produce a minimum transition Reynolds number on a 43-in. 10° cone. The results are an extension of previous work on an adiabatic cone and show a linear effect of wall temperature reduction on size of roughness. The dependence of roughness size on trip position, Mach number, and surface temperature is expressed in a simple relation.

Patent
Robert I Garrett1
20 May 1968
TL;DR: In this article, a planar bend is used to separate one or more components from a MULTICOMPONENT, high-pressure GAS STREAM and then the separable components are collected along with the DISSOLVED and ENTRAINED components.
Abstract: METHOD AND APPARATUS FOR SEPARATING ONE OR MORE COMPONENTS FROM A MULTICOMPONENT, HIGH-PRESSURE GAS STREAM THE GAS STREAM IS EXPANDED TO SUPERSONIC VELOCITY THROUGH A SUPERSONIC EFFUSER TO ACHIEVE LOW TEMPERATURES AND LOW PRESSURES IN THE SUPERSONIC GAS STREAM AND CAUSE CONDENSED LIQUID PARTICLES (DROPS) AND/OR SOLID PARTICLES TO FORM. THE SUPERSONIC GAS STREAM IS MADE TO TRAVERSE A PLANAR BEND PROVIDED WITH A PERMEABLE OUTER WALL TO AND THROUGH WHICH LIQUID AND/OR SOLID PARTICLES ARE INERTIALLY MOVED AND THEREBY SEPARATED FROM THE GAS STREAM. THE SEPARATED PARTICLES ARE COLLECTED ALONG WITH THE DISSOLVED AND ENTRAINED GASES WHICH ALSO SEPARATE FROM THE GAS STREAM. THE SUPERSONIC GAS STREAM IS THEN DECELERATED TO SUBSONIC FLOW THROUGH A SUPERSONIC DIFFUSER AND PART OF THE PRESSURE OF THE GAS STREAM IS RECOVERED. MEANS ARE PROVIDED TO MOVE THE FINAL SHOCK WAVE TO A STABLE POSITION WHICH IS AN OPTIMUM POSITION FOR PRACTICAL OPERATION WHEN A SUPERSONIC FLOW IS STARTED THROUGH THE SUPERSONIC FLOW SEPARATOR. TO ACHEIVE SUCH MOVEMENT OF THE FINAL SHOCK WAVE, THE THROAT AREA AND CONTOUR OF THE DIFFUSER IS MADE ADJUSTABLE SO THAT THE THROAT AREA AND SIZE OF THE DIFFUSER CHANNEL ARE INITIALLY ENLARGED TO MOVE THE FINAL SHOCK WAVE THROUGH THE DIFFUSER THROAT TO START SUPERSONIC FLOW IN THE SEPARATOR AND THEREAFTER REDUCED TO LOCATE THE FINAL SHOCK WAVE NEAR THE DIFFUSER THOAT TO OPERATE THE SEPARATOR EFFICIENTLY AT MAXIMUM BACK PRESSURE.

Journal ArticleDOI
TL;DR: In this article, separated flow in two dimensional and axisymmetric nozzles with various wall contours, determining flow structure and shock-boundary layer interaction is discussed.
Abstract: Separated flow in two dimensional and axisymmetric nozzles with various wall contours, determining flow structure and shock-boundary layer interaction

01 Mar 1968
TL;DR: In this paper, the effect of boattail juncture shape on pressure drag coefficients of isolated afterbodies of supersonic engine nacelles was investigated in the case of a single-passenger aircraft.
Abstract: Effect of boattail juncture shape on pressure drag coefficients of isolated afterbodies of supersonic engine nacelles

Proceedings ArticleDOI
01 Jan 1968
TL;DR: In this paper, the aerodynamic influence coefficients method was used for wing-body combinations analysis and design at supersonic and subsonic speeds by aerodynamic impact coefficients method.
Abstract: Wing-body combinations analysis and design at supersonic and subsonic speeds by aerodynamic influence coefficients method


Journal ArticleDOI
TL;DR: In this article, a review of different schemes of the numerical method of characteristics for calculating three-dimensional steady supersonic gas flow about bodies moving at incidence is given, and some results of numerical solutions are given for illustration.

01 Mar 1968
TL;DR: In this paper, boundary-layer transition and aerodynamic noise measurements were made on sharp leading-edge two-dimensional models in supersonic wind tunnels (1 ft to 16 ft).
Abstract: : Boundary-layer transition and aerodynamic noise measurements were made on sharp leading-edge two-dimensional models in supersonic wind tunnels (1 ft to 16 ft). These data showed, conclusively, a significant and continuous increase in transition Reynolds number and a significant decrease in radiated aerodynamic noise (generated by the tunnel wall turbulent boundary layer) with increasing tunnel size. Results obtained in the AEDC-VKF Tunnel A from a shroud configuration placed concentrically around a hollow cylinder transition model and a flat plate microphone model further demonstrated the strong adverse effects that radiated aerodynamic noise will have on transition. A correlation of transition Reynolds numbers based on transition data from nine wind tunnels (3 or = M or = 8) was developed. The correlation was dependent only on the tunnel wall, turbulent boundary layer, aerodynamic noise parameters (displacement thickness and skin friction), and the tunnel test section circumference.

01 Aug 1968
TL;DR: In this article, a Mach box digital computer program, MBOX, was developed and written for surfaces with trailing edge control surfaces, which is a very versatile and reliable method for obtaining steady and unsteady supersonic aerodynamic coefficients.
Abstract: : A Mach box digital computer program, MBOX, has been developed and written for surfaces with trailing edge control surfaces. It is a very versatile and reliable method for obtaining steady and unsteady supersonic aerodynamic coefficients that may be used in response and flutter analyses of rather general lifting surfaces. Wing configurations can include those with folded tips, cranked leading and trailing edges, and supersonic or subsonic leading and trailing edges. At the option of the user, MBOX will also calculate steady or unsteady lifting pressure distributions, and if the generalized mass and stiffness matrices are provided, it will obtain solutions of flutter equations in one computer run. (Author)


Journal ArticleDOI
TL;DR: In this paper, a massively blowing wedge was studied in a supersonic arc tunnel and the shock wave and separation streamline were visualized in the low-density flow by two chemiluminescent techniques.
Abstract: A massively blowing wedge was studied in a supersonic arc tunnel. The shock wave and separation streamline were visualized in the low-density flow by two chemiluminescent techniques. The data obtained under a variety of conditions were found to be correlated by a linear relation between two nondimensional blowing parameters derived by an approximate momentum balance analysis. Viscous interaction effects were found to be important in determining pressure levels and shock-wave angles at a Reynolds number of ^178. The relation between the blowing parameters found for the wedge also seems to hold for a blowing cone when the three-dimensional geometry is taken into account.



Patent
25 Apr 1968
TL;DR: In this article, reflecting surfaces are provided in the area of the maximum pressure difference of the Mach cone created by the aircraft, which reduces the supersonic boom caused by aircraft.
Abstract: The invention describes means for reducing the supersonic boom caused by aircraft. Underneath the aircraft reflecting surfaces are provided in the area of the maximum pressure difference of the Mach cone created by the aircraft.

Journal ArticleDOI
TL;DR: In this article, the authors investigated the pressure distribution due to the expansion of a boundary layer at a sharp trailing edge, where the entire expansion is assumed to be inviscid, and the subsonic and supersonic regions of the boundary layer were treated with the streamtube approximation and Prandtl-Meyer relation, respectively.
Abstract: The pressure distribution due to the expansion of a boundary layer at a sharp trailing edge is investigated analytically. The entire expansion is assumed to be inviscid, and the subsonic and supersonic regions of the boundary layer are treated with the streamtube approximation and Prandtl-Meyer relation, respectively. Matching of streamfunction and pressure on the initially sonic streamline determines the expansion process upstream of the trailing edge, whose location is uniquely determined. The extent of the upstream expansion increases linearly with the Mach number at the edge of the boundary layer for Mach numbers greater than five. At the trailing edge the wall pressure is less than its initial value, and the final expansion to the base pressure requires a smaller turning angle of the dividing streamline than was previously calculated. The effects of rotationality in the supersonic region of the boundary layer and the influence of viscosity are also considered.