Showing papers on "Supersonic speed published in 1982"

Broadband Shock Noise from Supersonic Jets

Abstract: Broadband shock noise from supersonic jets is investigated through acoustic measurements in both the near and far fields. The peak Helmholtz number of broadband shock noise from unheated convergent nozzles is found to be independent of nozzle pressure ratio when based on the length of the shock cells and the ambient speed of sound. Excellent agreement between power spectral densities measured at various far-field angles is obtained at and above the peak shock noise frequency when source convection and directivity effects are included. The directivity of broadband shock noise is found to be pointed in the upstream direction, with omnidirectionality being approached only at high pressure ratios. For both convergent and convergent-divergent nozzles, the relative importance of shock noise with respect to jet-mixing noise is found to be maximum near the pressure ratio at which a Mach disk begins to form in the jet. Near-field measurements point to a limited portion of the shock cell system as the region of dominant broadband noise emission from a highly underexpanded convergent nozzle.

How isolated are molecules in a molecular beam

Abstract: General expressions are developed for the number of collisions per second a molecule makes with other molecules in an effusive molecular beam, a supersonic molecule beam, and a supersonic jet. When characterized by the same density and the same local temperature parameter, we find that the collision frequencies in an effusive beam, supersonic beam, supersonic jet, and a bulb are approximately in the ratio 1:(3 /2) :3:3 . Collision rates appropriate to practical applications are discussed.

Measurements of Mean Static Pressure and Far Field Acoustics of Shock Containing Supersonic Jets

Abstract: The far field acoustic data base generated in studies of broadband shock noise from supersonic jets is presented. Both conical and contoured nozzles of exit Mach numbers 1.0, 1.5, and 2.0 were tested using unheated air at pressure ratios ranging from 1.9 to 14. Tests were performed both with and without screech suppression tabs. Overall sound pressure variations and representative 1/3-octave and narrowband spectra are presented. The mean static pressure measured within these jets is also surveyed.

Device for diagnosing body cavity interiors with supersonic waves

Abstract: A device for diagnosing the interior of a body cavity is described which includes an optical endoscope, and a supersonic wave scanner probe which may be physically and operatively attached thereto. The supersonic wave scanner may be used with the optical endoscope, or may be detached and replaced by a hemispherical lid whereupon the endoscope may be inserted into the body cavity for optical observations only.

Nonlinear stability and instability of transonic flows through a nozzle

Abstract: We study transonic flows along a nozzle based on a one-dimensional model. It is shown that flows along the expanding portion of the nozzle are stable. On the other hand, flows with standing shock waves along a contracting duct are dynamically unstable. This was conjectured by the author based on the study of noninteracting wave patterns. The author had shown earlier that supersonic and subsonic flows along a duct with various cross sections are stable. Basic to our analysis are estimates showing that shock waves tend to decelerate along an expanding duct and accelerate along a contracting duct.

Proton temperature anisotropy in the polar wind

Abstract: The steady state flow of a fully ionized H(+)-O(+) electron plasma along geomagnetic field lines in the high-latitude topside ionosphere is studied. The theoretical formulation is based on a 13-moment system of transport equations, and allows for different species temperatures parallel and perpendicular to the geomagnetic field and nonclassical heat flows. For subsonic and supersonic flows, an appreciable H(+) temperature anisotropy occurs at all altitudes above 1500 km, and tends to be regulated at high altitudes. The direction of the temperature anisotropy is related to the direction of the H(+) heat flow to some extent; for supersonic flow an upward flow of heat from the lower ionosphere is required, while for subsonic flow solutions can be obtained with a downward H(+) heat flow. For subsonic flow, H(+)-H(+) collisions have an important effect on the H(+) stress and heat flow balance at all altitudes between 1500 and 12,000 km.

Water-Tunnel Studies of Leading-Edge Vortices

Abstract: Flow visualization studies have been made at Northrop in a hydrodynamic facility of leading-edge vortex flows. Vortex core trajectory and stability characteristics have been obtained on wing planforms suitable for subsonic-transonic and supersonic cruise fighter designs. The applicability of results obtained at low Reynolds number in water to higher Reynolds number vortex flow phenomena in air is addressed. Comparisons of watertunnel vortex positions and burst locations are made with flow visualization results obtained in air. Where appropriate, correlation of water-tunnel vortex flow behavior is made with trends observed in subsonic windtunnel data.

CARS spectroscopy of molecules in supersonic free jets

Abstract: CARS spectra are reported for N2, O2, H2, D2, C2H4, and various mixtures cooled by supersonic jet expansion. Analysis of the Q branch structure of the diatomic molecules permits direct measurement of rotational populations, temperatures, and densities as a function of position in the free jet zone. Good agreement is obtained with temperature results for N2 from electron beam fluorescence studies and with densities calculated from the customary isentropic expansion equations. The detection sensitivity was sufficient that, for the first time, broadband CARS spectra were obtained for a molecular beam (D2). A simple relaxation model of the expansion process yields rotational collision numbers of 10±4 And 14±4 for N2 and O2 but these values may be high because of condensation heating. The model was not successful in treating the rotational relaxation of D2 and H2 but temperature measurements show qualitatively that D2 ⋅⋅⋅ X transfer efficiency increases in the order X = D2

Low-frequency sound radiation and generation due to the interaction of unsteady flow with a jet pipe

Abstract: In this paper we examine the low-frequency sound radiated when various types of unsteady flow interact with a jet pipe. In each case we solve the problem exactly by the Wiener-Hopf technique, producing results valid for arbitrary internal and ex- ternal Mach numbers and temperatures, discuss the importance of a Kutta condition at the duct exit, and provide an interpretation, in elementary terms, of the radiated sound field using the Lighthill acoustic analogy. A central feature is that the solutions are always obtained subject to a causality requirement, regardless of whether or not a Kutta condition is imposed at the pipe lip. When low-frequency sound propagates down the jet pipe, little of it reaches the far field, and the major disturbance outside the pipe is that associated with the jet instability waves. At subsonic jet speeds and low-enough Strouhal number these waves transport kinetic energy at a rate precisely balancing the loss of acoustic energy from the pipe, resulting in a net attenuation of the sound power. For supersonic jet condi- tions a further wave motion, the unsteady-flow counterpart of the steady wave struc- ture of an imperfectly expanded jet, is present in addition to the instability wave. We use the Lighthill acoustic analogy to show that, for high-enough jet Mach number and temperature, the sound radiation is caused largely by quadrupole sources arising from the jet instability waves. An alternative interpretation uses the acoustic analogy incorporating a mean flow due to Dowling, Ffowcs Williams and Goldstein, and expresses the far-field sound as the sum of contributions from monopoles and dipoles distributed over the duct exit. The directivity and power of the calculated far-field sound are in good agreement with experiments. We also calculate the sound scattered by the jet pipe when there is an incident external sound field, and show a previously published result to be in error. In general, the flbw phenomena produced by internal and external incident sound fields are similar. Finally, we discuss the effects of nozzle contraction. We find that the radiated sound field is little changed in character, but that the reflection properties of the nozzle may be drastically altered.

Supersonic jet noise generated by large scale instabilities

Abstract: The role of large scale wavelike structures as the major mechanism for supersonic jet noise emission is examined With the use of aerodynamic and acoustic data for low Reynolds number, supersonic jets at and below 70 thousand comparisons are made with flow fluctuation and acoustic measurements in high Reynolds number, supersonic jets These comparisons show that a similar physical mechanism governs the generation of sound emitted in he principal noise direction These experimental data are further compared with a linear instability theory whose prediction for the axial location of peak wave amplitude agrees satisfactorily with measured phased averaged flow fluctuation data in the low Reynolds number jets The agreement between theory and experiment in the high Reynolds number flow differs as to the axial location for peak flow fluctuations and predicts an apparent origin for sound emission far upstream of the measured acoustic data

A nonintrusive laser interferometer method for measurement of skin friction

Abstract: A method is described for monitoring the changing thickness of a thin oil film subject to an aerodynamic shear stress using two focused laser beams. The measurement is then simply analyzed in terms of the surface skin friction of the flow. The analysis includes the effects of arbitrarily large pressure and skinfriction gradients, gravity, and time-varying oil temperature. It may also be applied to three-dimensional flows with unknown direction. Applications are presented for a variety of flows including two-dimensional flows, three-dimensional swirling flows, separated flows, supersonic high-Reynolds-number flows, and delta-wing vortical flows.

Three-Dimensional Supersonic Viscous Flow over a Cone at Incidence

Abstract: : The viscous supersonic flow over a sharp cone at incidence is examined numerically with a coupled strongly implicit algorithm for the properties in the plane normal to the cone axis. The Navier-Stokes equations are considered in a boundary layer-like or parabolized manner and global relaxation is considered for the pressure interaction. It is shown that departure effects can be effectively eliminated by forward differencing for the axial pressure gradient. Moreover, this approximation retains the implicit free pressure interaction required for geometries where axial flow separation is possible. (Author)

Aerodynamic levitation of laser-heated solids in gas jets

Abstract: The aerodynamic levitation technique is developed for studies of high-temperature material properties and gas/condensed-phase reaction kinetics. Stable levitation is demonstrated in a supersonic jet from a 0.081 cm nozzle with 0.03-0.20 g 0.24-0.47 cm diameter solid spheres at a height between 0.7-2.0 cm above the nozzle and ambient pressures between 1.1-18 Torr. A model of supersonic jet levitation is developed which accurately predicts the values of height vs pressure over the full range of conditions investigated. It is found that the efficiency with which jet momentum is converted into levitation force decreases with the jet/specimen diameter ratio and the jet Reynolds number, and the rate of jet spreading with distance from the nozzle is found to agree with that measured by pitot tube traverses of the jet. In addition, laser heating is shown to reduce the jet momentum required for levitation at a given height and to increase levitation stability. Measurements of sphere levitation in subsonic gas jets show that the required jet momentum flow rate exceeds the specimen weight by about 2/the specimen drag coefficient at its terminal free-fall speed under ambient conditions.

Investigation of Supersonic Air Ejectors : Part 2,Effects of Throat-Area-Ratio on Ejector Performance

Abstract: The performance of a supersonic air ejector was experimentally studied over a range of ejector-to-nozzle throat-area-ratios Ψ and primary nozzle Mach numbers Mlp. The investigations covered both straight-tube and second-throat ejector types. As the result, the effects of the ratio Ψ on the ejector performance were made clear. The performance curves were classified into five groups by the values of Ψ and a primary chamber-to-ambient pressure ratio pop/pa. For a fixed Ψ, an optimum pop/pa exists which realizes a maximum secondary flow rate and a minimum secondary chamber pressure. Also, for a fixed Mlp, the optimum pop/pa becomes minimum at a certain Ψ, which was defined as an optimum throat-area-ratio. The physical meanings of the optimum pressure ratio and the optimum throat-area-ratio were clarified using the results of pressure measurements and optical observations.

Numerical Solution of Supersonic Viscous Flow over Blunt Delta Wings

Abstract: A general parabolized Navier-Stokes code has been developed to compute the steady supersonic viscous flow around arbitrary body shapes at high angles of attack. A nonorthogonal three-dimensional coordinate frame permits the code to march with solution surfaces which are the most appropriate. The code has been used to calculate the laminar flow over a slab delta wing with 70 deg sweep at angles of attack up to 41.5 deg and Mach numbers of 6.8 and 9.6. The computed shock shapes, surface pressures and heat transfer coefficients are compared with experiment and show agreement.

Caster type empennage assembly for aircraft

Abstract: An empennage assembly for supersonic aircraft includes longitudinally and rearwardly extending booms (12) mounted upon the wings (16) or fuselage of the aircraft. The booms (12) include aft rotatable sections (22) upon which are mounted larger (18) and smaller (20) tail surfaces. The boom sections (22) are angularly rotated through angular displacements θ such that the dispositions of the tail surfaces (18, 20) are interchanged between their dispositions during low subsonic and high supersonic flight conditions. In this manner, the directional and longitudinal aerodynamic static stability components of the aircraft are rendered substantially constant at an optimum low or near-neutral level of stability in order to enhance the flight maneuverability capabilities of the aircraft throughout the subsonic, transonic, and high supersonic speed ranges. The empennage (10) is also advantageously employed for enhancing the directional stability characteristics of all aircraft under varied angle of attack conditions, and still further, can likewise enhance the lift characteristics of STOL aircraft employing vectored thrust. In connection with the use of the present invention empennage system upon supersonic aircraft, the rotational orientation of the empennage system (10) and its associated boom sections (22) is automatically programmed by suitable feed-back control means (23) in response to sensed changes in Mach number. In a similar manner, the rotational orientation of the empennage system (10) and its associated boom sections (22) would also be automatically programmed by the feed-back control means (23) as a function of angle of attack.

Shockless Design and Analysis of Transonic Cascade Shapes

Abstract: A fast computer program was developed to eliminate the shocks by slightly altering portions of the contour of a given airfoil in the cascade. The program can be used in two basic modes: (1) An analysis for steady, transonic, potential flow through a given planar cascade of airfoils and (2) a design for converting a given cascade into a shockless transonic cascade. The design mode can automatically be followed by the analysis mode, which confirms that the flow field is shock free. The program generates its own multilevel boundary conforming computational grids and solves a full potential equation in a fully conservative form. The shockless design is performed by implementing Sobieczky's fictitious-gas elliptic continuation concept.

A Strongly Implicit Procedure for Steady Three-Dimensional Transonic Potential Flows

Abstract: A rapid relaxation algorithm for the numerical solution of the conservative full potential equation is described and used to predict transonic potential flow past thick swept wings. A sheared parabolic coordinate system is used to simplify the treatment of boundary conditions and an upwind bias is applied to the density in supersonic regions. A unique feature of the present work is the extension of a strongly implicit procedure, that has, in the past, been applied only to elliptic equations, to the present elliptic-hyperbolic problem. Numerical results are presented to demonstrate that the present method is considerably faster than existing line relaxation procedures.

Navier-Stokes Solutions for an Axisymmetric Nozzle

Abstract: Numerical solutions of the Navier-Stokes equations are obtained for an axisymmetric nozzle in a supersonic external flowfield (M^ =1.94, My =3.0, Re^ =2.2xl0). Five jet pressure ratio conditions ranging from a highly overexpanded case that exhibits a Mach disk shock formation to a slightly underexpanded case are solved computationally. MacCormack's explicit finite-difference algorithm, an adaptive grid scheme, and locally dependent eddy viscosity modeling are utilized to obtain the numerical solutions. The computational results accurately reproduce the experimentally observed viscous effects on the nozzle base pressure and shock locations resulting from a thick base annulus on the nozzle. Correct transition from regular shock reflection to Mach disk reflection in the jet core was achieved numerically.

Mean flow and noise measurements in a Mach 3.5 pilot quiet tunnel

Abstract: The use of Mach 3.5 two-dimensional rapid expansion nozzle for wind tunnel testing at supersonic speeds and low noise conditions encountered in high altitude flights is described. The supersonic pilot quiet tunnel is located at the NASA Langley Research Center and a description of the facility is provided, along with instrumentation and noise measurement test data at 30, 50, and 75 psia. The mean pitot pressure distributions, rms noise levels, the effect of unit Reynolds number, wall waviness, wall contaminants, and the effects of closing the bleed valve are analyzed. Typical laminar and turbulent spectra are presented, along with a summary of the effect of slot throat adjustment on the power spectra. Comparisons are made of the power spectra with the bleed valve open and closed, and of the rms fluctuating pressures with levels from conventional nozzles, and the performance capabilities are evaluated for use in transition studies.