Showing papers on "Freestream published in 1981"

Improved Calculation of High Speed Inlet Flows. Part II: Results

Abstract: An improved numerical algorithm that solves the full mean compressible Navier-Stokes equations has been applied to the calculation of the flowfield in three separate configurations of a simulated high speed aircraft inlet. The inlet geometry consists of a converging supersonic diffuser, formed by two nonparallel plates, followed by a constant height "throat." For all cases, the freestream Mach number is 3.51, and the Reynolds number is 13.6 X10 based on the inlet length. The three configurations are characterized by different values of the angle of the converging supersonic diffuser and different boundary-layer bleed schedules. The computed results are compared with detailed experimental data for the ramp and cowl surface pressure distribution, and the boundary-layer pitot profiles at seven different streamwise locations. The agreement with the experimental results is generally good, although the experimental data display evidence of three dimensionality over a portion of the inlet flowfield.

Optimum tail shapes for bodies of revolution

Abstract: This paper has n dual purpose: to describe a method of designifig short tails for bodies 01 revolution that sutisfy Stratford's criterion for zero shear at the wall, and to show n few shapes that have been calculated. Stratford's original two-dimensional solution, extended to axisymmetric flow, has been used to implement the prucedure. The method involves simultrmneous solution of the extended Stcatford equation together with the necessary boundary conditions by means of an inverse potential flow program. Tails designed by this procedure are entirely at incipient separmtion (no skin friction); therefore the pressure recovery Is the most rapid possible, making the resultant tail the shortest possible, subject to no separation. The Final result Is a geometry uniquely determined for freestream conditions, the transition point, and of course the basic forebody. The computer program can operate in one of two modes: 1) the forebody geometry can be malntained (except for a small region near the tall juncture) with only the tail shape determined by the method or 2) the forebudy ve1ocity distribution can he malntalned up to the paint of pressure recovery. The forebody geometry wlll then be altered for some distance upstream of the tail juncture. A number of solutions are presented for both of the above modes. Nomenclature A = reference area CIIYOj =drag coefficient based on (vol~me)~ C,,* =drag coefficient based on frontal area C, =pressure coefficient, C, = I - u2 /u&C, = Stratford type pressure coefficient C, = 1 - u2 /ud, I = reference length L = length representative of the length of the body r =radius of body at any point t?, = Reynolds number, U,S/U s =distance along body surface, see Fig. I u =velocity along body outside the boundary layer u, =freestreamvelocity SIA bscripts

Determination of atmospheric properties for STS-1 aerothermodynamic investigations

Abstract: A procedure for determining an approximation to the freestream atmospheric properties along the Shuttle entry trajectory is presented. Meteorological data as input is obtained by rawinsondes from surface to 70 km, and meteorological spheres from 60-90 km, launched from Hawaii and California. The Langley Atmospheric Information Retrieval System (LAIRS) developed to approximate the atmospheric freestream properties along the flight path, is outlined, noting temperature and wind data are interpolated in altitude, while gradients and diurnal and semidiurnal coefficients are taken from the COSPAR reference atmosphere. The data are input to a model to project temperature profiles for the Shuttle descent, and the input atmospheric parameters are listed. Efforts are continuing in order to correct discrepancies in the generated profiles for regions below 3 km.

Computation of airfoil buffet boundaries

Abstract: The ILLIAC IV computer has been programmed with an implicit, finite-difference code for solving the thin layer compressible Navier-Stokes equation. Results presented for the case of the buffet boundaries of a conventional and a supercritical airfoil section at high Reynolds numbers are found to be in agreement with experimentally determined buffet boundaries, especially at the higher freestream Mach numbers and lower lift coefficients where the onset of unsteady flows is associated with shock wave-induced boundary layer separation.

Mono-element combined supercritical high lift airfoil

Abstract: A multi-purpose mono-element airfoil is disclosed for aerodynamic and hydrodynamic vehicles and devices. The multipurpose mono-element highlift airfoil when utilized in an aerodynamic application provides a combined no-moving-parts high lift and cruise airfoil which in conjunction with a plenum, upon pressure initiation, causes pressurized air to issue from a slot tangent to the airfoil surface and remains attached to the airfoil's shaped trailing edge, providing a controlled resultant force of thrust. Upon application to hydrodynamic vehicles, the multi-purpose mono-element airfoil is placed in the freestream and provides turning or pitching forces to the vehicle without any deflection of itself or any mechanical components.

Aerodynamic Interference Between Two Darrieus Wind Turbines

Abstract: The effect of aerodynamic interference on the performance of two curved bladed Darrieus-type vertical axis wind turbines has been calculated using a vortex/lifting line aerodynamic model. The turbines have a tower-to-tower separation distance of 1.5 turbine diameters, with the line of turbine centers varying with respect to the ambient wind direction. The effects of freestream turbulence were neglected. For the cases examined, the calculations showed that the downwind turbine power decrement (1) was significant only when the line of turbine centers was coincident with the ambient wind direction, (2) increased with increasing tipspeed ratio, and (3) is due more to induced flow angularities downstream than to speed deficits near the downstream turbine.

Correlation of artificially induced boundary-layer transition data at hypersonic speeds

Abstract: Previous investigations of distributed-r oughness boundary-layer trips indicated that they are superior to spherical-type trips in that equally effective distributed-roughness trips are one-fifth as high and produce substantially smaller flowfield disturbances. The present work has expanded the data base, permitting correlation of distributed-roughness tripping data. The correlation thus developed includes a wide range of Reynolds numbers, cone angles, and trip heights. Plots are provided that permit the selection of distributed-r oughness trips without the need of boundary-laye r solutions. Nomenclature k = trip element height, in. M^ = freestream Mach number PE = pressure at the end of the roughness area (s/r,, =5), psia P'0 = freestream pitot pressure, psia p^ = freestream pressure, psia q^ = freestream dynamic pressure, psia ReeQ = Reynolds number based on boundary-layer edge conditions and momentum thickness Reer — Reynolds number based on boundary-laye r edge conditions and model nose radius Re^/ft = Reynolds number based on freestream conditions and a 1 ft length Re^>r = Reynolds number based on freestream conditions and model nose radius rb = model base radius, in.

Stagnation region gas film cooling: Spanwise angled injection from multiple rows of holes

Abstract: The stagnation region of a cylinder in a cross flow was used in experiments conducted with both a single row and multiple rows of spanwise angled (25 deg) coolant holes for a range of the coolant blowing ratio with a freestream to wall temperature ratio approximately equal to 1.7 and R(eD) = 90,000. Data from local heat flux measurements are presented for injection from a single row located at 5 deg, 22.9 deg, 40.8 deg, 58.7 deg from stagnation using a hole spacing ratio of S/d(o) = 5 and 10. Three multiple row configurations were also investigated. Data are presented for a uniform blowing distribution and for a nonuniform blowing distribution simulating a plenum supply. The data for local Stanton Number reduction demonstrated a lack of lateral spreading by the coolant jets. Heat flux levels larger than those without film cooling were observed directly behind the coolant holes as the blowing ratio exceeded a particular value. The data were spanwise averaged to illustrate the influence of injection location, blowing ratio and hole spacing. The large values of blowing ratio for the blowing distribution simulating a plenum supply resulted in heat flux levels behind the holes in excess of the values without film cooling. An increase in freestream turbulence intensity from 4.4 to 9.5 percent had a negligible effect on the film cooling performance.

Harmonic optimization of a periodic flow wind tunnel

Abstract: This work describes a wind-tunnel modification designed to superpose on the mean velocity sinusoidal longitudinal velocity fluctuations with minimal harmonic content. The technique is presented in light of a theoretical analysis of the low-frequency performance illustrating how harmonic suppression can be achieved with this particular design. Velocity fluctuations are produced by a system of primary rotating vanes and a bypass containing a secondary set of rotating vanes. Experimental data on tunnel performance are also presented. A significant reduction of the second harmonic content of the freestream velocity oscillations was achieved by adjustment of the bypass flow.

Freestream turbulence effects on the development of a rotor wake

Abstract: The effect of freestream turbulence on the development of a three-dimensional wake of a compressor rotor blade was studied experimentally. The turbulence level at the inlet of a rotor was varied systematically using grids upstream of the rotor. The rotor wake was measured with inlet turbulence intensities of 0.5, 3, and 5%. The experimental results indicate that the maxium change in the mean velocity defect is 4% over the range of inlet turbulence levels employed, while the turbulence structure in the wake is altered more substantially. The freestream turbulence effect was also analyzed, numerically, using the modified Reynolds stress closure model. The comparison between numerical prediction and experimental data shows that the freestream turbulence effect can be represented successfully with the turbulence closure model employed in this paper.

Autoignition characteristics of no. 2 diesel fuel

Abstract: Parametric tests to map the ignition delay characteristics were conducted at pressures of 3, 4, and 5 atm, inlet air temperatures up to 1150 K and fuel air equivalence ratios ranging from 02 to 10 Ignition delay times in the range of 6 msec to 60 msec at freestream flow velocities ranging from 10 m/sec to 40 m/sec were obtained The ignition delay times appeared to correlate with the inverse of pressure and the inverse exponent of temperature

Experimental investigation of the asymmetric body vortex wake

Abstract: An experimental investigation of the asymmetric body vortex wake of a circular cylinder in high subsonic flow is presented. Laser velocimeter, force and moment, and surface hot wire measurements were obtained for a freestream Mach number of 0.6 and Reynolds number (based on body diameter) of 0.62 x 10 to the 6th. Two component laser velocimeter measurements were made at three body cross-flow planes, x/d = 4, 8, and 12, and angles of attack of 25, 35, and 45 deg. Laser vapor screen photographs were also obtained at these body stations and angles of attack. Surface hot wire measurements were used to determine if any vortex switching occurred at various angles of attack of the body. The laser velocimeter measurements are related to the vapor screen photographs and side force measurements. These results show that more than one asymmetric body vortex wake configuration can exist for the same angle of attack and body roll angle.

Utilization of an Electron Beam for Density Measurements in Hypersonic Helium Flow

Abstract: S methods have been investigated for making nondestructive measurements of various wind tunnel parameters. Herein is described a study to determine the feasibility of using an electron beam to measure the freestream static density of gaseous helium over a range of hypersonic flow conditions. Measurements were made for a range of stagnation pressures and temperatures which produced freestream number densities of 1.53xl0 to 1.25x IO molecules/m and static temperatures from 2 to 80 K. The results showed the collision quenching cross section to be 4.4 x 10 ~ cm at 1 K and to have a weak temperature dependence of T'. Knowing these values, the freestream number density can be determined quite accurately. These results are reported in more detail in Ref. 1.

Compressibility Effects on Cavitation in High Speed Liquid Flow : First Report-Subsonic Liquid Flow

Abstract: Firstly the freestream Mach number for subsonic liquid flow is defined as being smaller than the lower critical Mach number in which the critical condition can be found on the cavity streamline The governing equation of elliptic type for velocity potential is solved by local linearization technique and then the corresponding relation are derived between subsonic and incompressible liquid f1ows Compressibi1ity effects are minutely discussed on the cavity characteristics of symmetrical wedge through the Mach number range and also comparisons are made with the former linearized analysis

Numerical study of re-entry vehicle aerodynamics in steady coning motion

Abstract: The objective of the present investigation is to formulate a numerical model to predict the aerodynamics of a hi-speed re-entry vehicle (RV). This analytical model is able to estimate the aerodynamic effects of the rotational motions (e.g., spinning and precession) of the vehicle as well as those effects resulting from its translational motions. Numerical results are presented for the three-dimensional flowfield about a blunt cone in a steady coning motion. The theoretical formulation consists of a two-layer model: the outer inviscid flow and the inner laminar and/or turbulent boundary layers. One significant contribution of the present paper is the estimation of the dynamic stability properties such as pitchand roll-damping derivatives. Effects of freestream Mach number, vehicle bluntness ratio, angle of attack, and the center of gravity location upon the flow properties are discussed.

Effect of Modification of the Trailing Edge of a Separating Wall on the Downstream Mixing of Parallel Flowing Streams.

Abstract: : The objective of this thesis is to study ways to enhance the mixing of two parallel streams of air by modifying the trailing edge of a separating wall. An apparatus was designed which achieved two-dimensional, good quality flow near the center of the test section passage, with freestream turbulence intensity of 2.2 percent. Measurements of the wake were made varying the velocity of one stream down to 37.5 percent of the other stream velocity, both in and upstream of the asymptotic region of the wake. A single element hot wire was used to measure velocity and RMS readings. The flat plate trailing edge was then slotted with five, eight millimeter slots and re-tested. The higher turbulence and wider wake of the flat plate indicate that the slotted plate wake does not achieve as good mixing as the flat plate wake. No velocity ratio of the two streams was found to maximize the wake growth for either configuration. Wake growth doubled when the slower velocity was 0.40 of the faster velocity. (Author)

An experimental investigation of a large delta P settling chamber for a supersonic pilot quiet tunnel

Abstract: The mean streamwise flow distributions and turbulence levels across the chamber were measured with a hot wire anemometer downstream of a series of porous Rigimesh plates which were shown to be an effective means of reducing the chamber acoustic disturbance levels due to upstream pipe and valve systems. Tests made with various types of flow conditioners downstream of the porous plates showed that a series of screens was the most effective means of achieving the objective of a uniform mean flow distribution with reduced vorticity levels downstream of the porous components. Frequency spectra obtained across the series of screens show that they reduce vorticity over a wide frequency range for several different initial upstream vorticity conditions. Improvements in the mechanical installation of the porous plates and damping screens and the use of porous plates with more uniform porosity should reduce the freestream velocity fluctuations to the minimum acoustic levels of about 0.5 percent.

Prediction of laminar-to-turbulent boundary layer transition in axial-flow turbomachinery

Abstract: : This paper first reviews current analytical studies regarding the laminar-to-turbulent boundary layer transition process, including relevant analysis of laminar separation, and then covers analytical and experimental studies concerning parametric effects on transition pertinent to axial-flow turbomachine blade rows. These include the effects of pressure gradient, freestream turbulence, surface roughness, heat transfer, and combined effects. Finally, the work of various authors who have established data correlations for transition on axial-flow turbomachine blade rows is evaluated. Their predictive equations and charts have been presented as a reference guide for determining the transition region on blades. Additionally, equations for calculating transitional boundary layer growth are presented. (Author)

An Analysis of the Inlet Plane Aerodynamics of an Oscillating Subsonic Cascade: An Evaluation of Experimental Data.

Abstract: : An analysis has been performed on unsteady pressure data measured in the leading edge region of a subsonic cascade oscillating in pitch. The objectives of the investigation were to determine the sensitivity of the freestream velocity disturbance at the inlet plane to variations in interblade phase angle, to compare the behavior of the freestream velocity disturbance as a function of interblade phase angle with that of both the inlet area oscillation and the measured moment response of the cascade, and to determine the influence of the gap-to-chord ratio on the blade loading near the leading edge. In this study, it was found that the behavior of the unsteady freestream flow entering an oscillating subsonic cascade is strongly influenced by the interblade phase angle. A more important finding was that the primary trends in the freestream disturbance match those of both the inlet area oscillation and the moment response of the cascade. Because of this correlation, a hypothesis is presented that describes how the interaction between the inlet area oscillation and the freestream flow might be relevant to the behavior of the cascade moment response and thus to the stability of the cascade motion. Analytical predictions based on this hypothesis are in qualitative agreement with the measured trends. The effect of gap-to-chord ratio on the unsteady blade loading was studied and found to be substantial in two ways: it alters the degree of influence of leading edge dynamic stall on the chordwise load distribution, and it alters the degree of influence of the interblade phase angle on the amplitude of the load response. (Author)