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Three-dimensional flow phenomena in a shrouded inducer were studied with a three-dimensional Navier-Stokes method. The details of the three-dimensional flow structure inside the inducer at design and off-design conditions are analyzed and the results are compared with some flow visualization results obtained at the California Institute of Technology.
A CFD solver has been developed to provide the time averaged details of the fountain flow typical for tiltrotor aircraft in hover. This Navier-Stokes solver, designated as ROTTILT, assumes the 3-D fountain flowfield to be steady and incompressible. The theoretical background is described in this manual. In order to enable the rotor trim solution in the presence of tiltrotor aircraft components such as wing, nacelle, and fuselage, the solver is coupled with a set of trim routines which are highly efficient in CPU and suitable for CFD analysis. The Cartesian grid technique utilized provides the user with a unique capability for insertion or elimination of any components of the bodies considered for a given tiltrotor aircraft configuration. The flowfield associated with either a semi or full-span configuration can be computed through user options in the ROTTILT input file. Full details associated with the numerical solution implemented in ROTTILT and assumptions are presented. A description of input surface mesh topology is provided in the appendices along with a listing of all preprocessor programs. Input variable definitions and default values are provided for the V22 aircraft. Limited predicted results using the coupled ROTTILT/WOPWOP program for the V22 in hover are made and compared with measurement. To visualize the V22 aircraft and predictions, a preprocessor graphics program GNU-PLOT3D was used. This program is described and example graphic results presented.
Results are presented of an experiment conducted to investigate potential sources of noise in the flow developed by two 22-in. diameter turbofan models. The R4 and M5 rotors that were tested were designed to operate at nominal take-off speeds of 12,657 and 14,064 RPMC, respectively. Both fans were tested with a common set of swept stators installed downstream of the rotors. Detailed measurements of the flows generated by the two were made using a laser Doppler velocimeter system. The wake flows generated by the two rotors are illustrated through a series of contour plots. These show that the two wake flows are quite different, especially in the tip region. These data are used to explain some of the differences in the rotor/stator interaction noise generated by the two fan stages. In addition to these wake data, measurements were also made in the R4 rotor blade passages. These results illustrate the tip flow development within the blade passages, its migration downstream, and (at high rotor speeds) its merging with the blade wake of the adjacent (following) blade. Data also depict the variation of this tip flow with tip clearance. Data obtained within the rotor blade passages at high rotational speeds illustrate the variation of the mean shock position across the different blade passages.
A wind tunnel test was conducted in the Langley Research Center 31-inch Continuous Flow Hypersonic Wind Tunnel from May 6, 1975 through June 3, 1975. The primary objectives of this test were the following: (1) to study the ability of the wind tunnel to repeat, on a run-to-run basis, data taken for identical configurations to determine if errors in repeatability could have a significant effect on jet interaction data, (2) to determine the effect of aerodynamic heating of the scale model on jet interaction, (3) to investigate the effects of elevon and body flap deflections on jet interaction, (4) to determine if the effects from jets fired separately along different axes can be added to equal the effects of the jets fired simultaneously (super position effects), (5) to study multiple jet effects, and (6) to investigate area ratio effects, i.e., the effect on jet interaction measurements of using wind tunnel nozzles with different area ratios in the same location. The model used in the test was a .010-scale model of the Space Shuttle Orbiter Configuration 3. The test was conducted at Mach 10.3 and a dynamic pressure of 150 psf. RCS chamber pressure was varied to simulate free flight dynamic pressures of 5, 7.5, 10, and 20 psf.
An approximate computational technique has been developed for predicting inviscid, radiating flows about blunt probes entering atmospheres consisting of hydrogen and helium. The technique is rapid and versatile and is well suited for performing parametric trade studies for outer planet entries. Details of the computational technique, the thermodynamic correlations, the 58-step absorption coefficient model and the analytic shock shape equations are discussed. Good comparisons of the radiative heating computed by the approximate method and by detailed calculations are obtained.
A Navier-Stokes calculation procedure is applied to high Reynolds number flow about steady and unsteady airfoils. The procedure solves the ensemble averaged governing equations via a linearized block implicit (LBI) technique which in general converges to a nominal steady flow within 120 time steps. The grid used in the computation is highly stretched thus resolving the turbulent boundary layers. Calculations have been compared with data for both NACA 4412 airfoil at high incidence and an NACA 0012 airfoil oscillating in dynamic stall. In both cases, good agreement is noted between measured and calculated surface pressure distributions.
A finite-volume lower-upper (LU) implicit scheme is used to simulate an inviscid flow in a tubine cascade. This approximate factorization scheme requires only the inversion of sparse lower and upper triangular matrices, which can be done efficiently without extensive storage. As an implicit scheme it allows a large time step to reach the steady state. An interactive grid generation program (TURBO), which is being developed, is used to generate grids. This program uses the control point form of algebraic grid generation which uses a sparse collection of control points from which the shape and position of coordinate curves can be adjusted. A distinct advantage of TURBO compared with other grid generation programs is that it allows the easy change of local mesh structure without affecting the grid outside the domain of independence. Sample grids are generated by TURBO for a compressor rotor blade and a turbine cascade. The turbine cascade flow is simulated by using the LU implicit scheme on the grid generated by TURBO.
A finite-volume lower-upper (LU) implicit scheme is used to simulate an inviscid flow in a turbine cascade. This approximate factorization scheme requires only the inversion of sparse lower and upper triangular matrices, which can be done efficiently without extensive storage. As an implicit scheme it allows a large time step to reach the steady state. An interactive grid generation program (TURBO), which is being developed, is used to generate grids. This program uses the control point form of algebraic grid generation which uses a sparse collection of control points from which the shape and position of coordinate curves can be adjusted. A distinct advantage of TURBO compared with other grid generation programs is that it allows the easy change of local mesh structure without affecting the grid outside the domain of dependence. Sample grids are generated by TURBO for a compressor rotor blade and a turbine cascade. The turbine cascade flow is simulated by using the LU implicit scheme on the grid generated by TURBO.
A discussion is given of the application of Fourier series techniques to the problems of linearized supersonic flow. The formulation presented is an extension of the doublet type of "fundamental solution" to higher order types of singularity. The equations developed have application to wing theory but are primarily of importance in wing-body interaction problems. A specific example of a wing-body interference problem is discussed in light of the presented methods.
Supersonic inviscid flow of equilibrium real gas over blunt axisymmetric bodies at large angles of attack including Apollo and Gemini shapes
A new, simple method is described for measuring and visualizing air flows. The method involves projecting a small heated metal pellet through the air at a speed greater than the flow. The pellet burns as it moves through the air and leaves a wake of very fine, visible, metal oxide particles. The position of this visible smoke trail is then photographed at a sequence of times. The displacement of the trail can be used to provide a plot of the normal component of velocity as a function of distance. Examples are given for very low speed thermal convection (less than about 1 m/sec) and low speed flow over airfoils and cylinders (less than about 10 m/sec). Comparisons of the method to pulsed smoke-wire, spark-tracer and laser fluorescence methods, which give similar information, are discussed.
This presentation will discuss Ir for online diagnostics with acoustics; Ptot, Ttot, Pstat rake surveys for mean flow measurements; focused Schleiren for density and some turbulence structure; Laser sheet visualization for near-nozzle diagnostics; and Two-point hotwire measurements for turbulence models.
The objectives of this project were to: (1) determine the effects of partial admission flow on rotor performance as a function of circumferential location and on unsteady rotor loading; and (2) Provide an efficient technique for determining turbine performance. Full admission simulation ws performed for the Simplex turbine and partial admission simulation is underway for the Simplex turbine.
In-flight flow visualization results of the vortical flow on the forebody and leading-edge extensions (LEX) of an F-18 research aircraft have been presented for angles of attack from 15.8 to 42.5 deg and for sideslip angles up to 7.5 deg. Water tunnel results using a 3-percent scale F-18 model and a variety of wind tunnel results are used for comparison and interpretation of the flight results. The LEX vortex core breakdown point moved forward with increasing angle of attack. For a constant angle of attack, the windward LEX vortex core breakdown moves forward and inboard with sideslip and the leeward vortex core breakdown moves aft and outboard. For a constant angle of attack, the windward location of interaction moved aft with increasing sideslip and the leeward interaction moved forward.
Higher power, high efficiency gas turbine engines require optimization of the seals and secondary flow systems as well as their impact on the powerstream. This work focuses on two aspects: 1. To apply the present day CFD tools (SCISEAL) to different real-life secondary flow applications from different original equipment manufacturers (OEM s) to provide feedback data and 2. Develop a computational methodology for coupled time-accurate simulation of the powerstream and secondary flow with emphasis on the interaction between the disk-cavity and rim seals flows with the powerstream (SCISEAL-MS-TURBO). One OEM simulation was of the Allison Engine Company T-56 turbine drum cavities including conjugate heat transfer with good agreement with data and provided design feedback information. Another was the GE aspirating seal where the 3-D CFD simulations played a major role in analysis and modification of that seal configuration. The second major objective, development of a coupled flow simulation capability was achieved by using two codes MS-TURBO for the powerstream and SCISEAL for the secondary flows with an interface coupling algorithm. The coupled code was tested against data from three differed configurations: 1. bladeless-rotor-stator-cavity turbine test rig, 2. UTRC high pressure turbine test rig, and, 3. the NASA Low-Speed-Air Compressor rig (LSAC) with results and limitations discussed herein.
Results are given of a comparative study of numerical procedures for computing solid wall boundary points in supersonic inviscid flow calculatons. Twenty five different calculation procedures were tested on two sample problems: a simple expansion wave and a simple compression (two-dimensional steady flow). A simple calculation procedure was developed. The merits and shortcomings of the various procedures are discussed, along with complications for three-dimensional and time-dependent flows.
Tests were conducted in the NASA Langley Research Center 31-inch Continuous Flow Hypersonic Wind Tunnel to determine RCS jet interaction effects on hypersonic aerodynamic characteristics and to investigate RT (gas constant times temperature) scaling effects on the RCS similitude. The model was an 0.010-scale replica of the Space Shuttle Orbiter Configuration 3. Hypersonic aerodynamic data were obtained from tests at Mach 10.3 and dynamic pressures of 200, 150, 125, and 100 psf. The RCS modes of pitch, yaw, and roll at free flight dynamic pressure simulation of 20 psf were investigated.
A description of the Axial Flow Turbine Research Facility (AFTRF) installed at the Turbomachinery Laboratory of the Pennsylvania State University is presented in this paper. The facility diameter is 91.66 cm (3 feet) and the hub-to-tip ratio of the blading is 0.73. The flow path consists of turbulence generating grid, 23 nozzle vane and 29 rotor blades followed by outlet guide vanes. The blading design, carried out by General Electric Company personnel, embody modern HP turbine design philosophy, loading and flow coefficient, reaction, aspect ratio, and blade turning angles; all within the current aircraft engine design turbine practice. State-of-the-art quasi-3D blade design techniques were used to design the vane and the blade shapes. The vanes and blades are heavily instrumented with fast response pressure, shear stress, and velocity probes and have provision for flow visualization and laser Doppler anemometer measurement. Furthermore, provision has been made for detailed nozzle wake, rotor wake and boundary layer surveys. A 150 channel slip ring unit is used for transmitting the rotor data to a stationary instrumentation system. All the design objectives have been met.