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A three-dimensional parabolic Navier-Stokes code, PEPSIG, was used to analyze the flow in the subsonic diffuser section of a typical modern inlet design. The effect of curvature of the diffuser centerline and transitioning cross sections was studied to determine the primary cause of flow distortion in the duct. Total pressure values at the engine compressor face are reported.
The cooling of gas turbine components was the subject of considerable research. The problem is difficult because the available coolant, compressor bleed air, is itself quite hot and has relatively poor thermophysical properties for a coolant. Injecting liquid water to evaporatively cool the air prior to its contact with the hot components was proposed and studied, particularly as a method of cooling for contingency power applications. Injection of a small quantity of cold liquid water into a relatively hot coolant air stream such that evaporation of the liquid is still in process when the coolant contacts the hot component was studied. No approach was found whereby heat transfer characteristics could be confidently predicted for such a case based solely on prior studies. It was not clear whether disequilibrium between phases at the inlet to the hot component section would improve cooling relative to that obtained where equilibrium was established prior to contact with the hot surface.
A numerical method is developed to analyze the inviscid flowfield of a high speed inlet by the solution of the Euler equations. The lower-upper implicit scheme in conjunction with adaptive dissipation proves to be an efficient and robust nonoscillatory shock capturing technique for high Mach number flows as well as for transonic flows.
Problems encountered in obtaining good engine face pressure data were studied. A single static measurement located upstream of the engine hub in the stream flow was found to provide a pressure signal suitable for engine control. Two identical probes for measuring fan inlet static (PS2) pressure were designed and mounted on the hub of the left F100-PW-100 turbofan engine installed in the F-15 test aircraft for flight evaluation. The probe is used as a static pressure sensor for a digital engine control system.
Electromagnetic backscattering from an open ended three dimensional inlet model is analyzed and computed patterns are compared with results of experimental measurements. The model is comprised of two sections. The first section consists of a linearly tapered waveguide with a rectangular opening at one end and the other end is connected to the second section which is a uniform rectangular waveguide with a planar perfectly conducting termination. The model is electrically large so that many propagating modes are excited. The method of analysis contains conventional aperture integration and modal techniques combined with high frequency techniques, which employ concepts such as modal rays, geometrical theory of diffraction and equivalent currents. For the cases considered, it is shown that only a few of the many propagating modes contribute appreciably to the backscattered field. These modes are selected according to their modal ray angle directions.
This paper describes numerical simulations of self-excited oscillations in a two-dimensional transonic inlet-diffuser flow by solving the Navier-Stokes equations with a two-equation turbulence model. The calculated amplitudes of oscillations for the terminal shock and the velocity fields compare well with experimental measurements; however, the predicted frequency of oscillations is about 50 percent higher. The formation of a pair of downstream-traveling, counter-rotating vortices at each cycle of velocity fluctuations, as reported experimentally, is vividly revealed by the numerical results.
The demand for increased gas turbine engine efficiency with minimum weight is leading to more complex blading designs, in which viscous and three-dimensional effects are significant. As a result, design procedures based on first principle, experimentally verified, three-dimensional aerodynamic analyses are required. This paper describes a series of experiments performed in a large-scale, subsonic, annular cascade facility specifically designed to provide three-dimensional aerodynamic data suitable for code verification. In particular, the effect of inlet velocity profile on the overall three-dimensional performance of a classical flat plate airfoil cascade is investigated over a range of incidence angles including those resulting in airfoil surface flow separation. All of the data are analyzed and correlated with appropriate nonseparated flow predictions.
A numerical method is developed to analyze the inviscid flowfield of a high speed inlet by the solution of the Euler equations. The LU implicit scheme in conjunction with adaptive dissipation proves to be an efficient and robust nonoscillatory shock capturing technique for high Mach number flows as well as for transonic flows.
NASCRIN program analyzes two-dimensional flow fields in supersoniccombustion ramjet (scramjet) inlets. Solves two-dimensional Euler or Navier-Stokes equations in conservative form by unsplit, explicit, two-step finite-difference method. More recent explicit/implicit, two-step scheme incorporated by analysis of viscous flow. Algebraic, two-layer eddy-viscosity model used for turbulent-flow calculations. Vectorized version, written for CDC CYBER 205, whereas scalar version, can be run on CRAY or other scalar computers.
A variable inlet guide van (VIGV) type convertible engine that could be used to power future high-speed rotorcraft was tested on an outdoor stand. The engine ran stably and smoothly in the turbofan, turboshaft, and dual (combined fan and shaft) power modes. In the turbofan mode with the VIGV open fuel consumption was comparable to that of a conventional turbofan engine. In the turboshaft mode with the VIGV closed fuel consumption was higher than that of present turboshaft engines because power was wasted in churning fan-tip airflow. In dynamic performance tests with a specially built digital engine control and using a waterbrake dynamometer for shaft load, the engine responded effectively to large steps in thrust command and shaft torque. Previous mission analyses of a conceptual X-wing rotorcraft capable of 400-knot cruise speed were revised to account for more fan-tip churning power loss than was originally estimated. The new calculations confirm that using convertible engines rather than separate lift and cruise engines would result in a smaller, lighter craft with lower fuel use and direct operating cost.
Calculations are presented predicting the onset of flow instability for a multistage low speed axial compressor operating in circumferentially distorted inlet flow. The most important feature of the model used is that it attempts to properly account for the fluid dynamic interaction between the spoiled and unspoiled sectors of the compressor. The calculations show that there is an approximate stability criterion, the annulus averaged slope of the compressor pressure rise characteristic equal to zero, that is valid whenever the dynamics of the compressor distorted flowfield can be considered independent of the compressor environment. This approximate criterion is used to investigate the relationship between the present model and the parallel compressor model. Further calculations are performed to investigate cases of interest when the dynamics of the flowfield are coupled to the environment. Resonant cases and cases when the distortion is unsteady are studied. In particular, it is shown that rotating distortions which propagate in the rotor direction can have a greater effect on stability margin than stationary or counter-rotational ones. Finally, it is shown that the general predictions of the model are insensitive to the details of the unsteady bladerow dynamics.
The paper outlines an approach to the design of three-dimensional inlets for scramjet engines. The basis of the techniques used is the method of streamline tracing through an inviscid axisymmetric flow field. A technique is described for making a smooth change of cross-section shape from rectangular to circular. A feature is the considerable use of computer-graphics to provide a 'user-oriented' procedure which can produce promising design configurations for subsequent analysis with CFD codes. An example is given to demonstrate the capabilities of the design techniques.
This paper describes the results of a program initiated to identify the cause and the mechanism of 4-kHz vibrations of the Space Shuttle main engine's (SSME's) LOX inlet tee, which were revealed in 17 percent of the SSMEs during ground tests. The results of the engines' inspection, structural analyses, computational fluid dynamics studies, literature reviews, expert consultations, and laboratory testings identified the probable cause of the 4-kHz vibrations being related to the fact that the walls of some tees were thinner than in the mojority of tees, resulting in less stiff structures. The mechanism of the vibrations was identified as vortex shedding in combination with a structural coupling with the thrust cone, which increased the vane mobility, which, in turn, aggravated the fluid coupling with the thrust cone. As a result, the vane mobility was increased, aggravating the fluid coupling with the vanes. An internal modification was developed and was successfully demonstrated on an engine that had previously exhibited a 4-kHz problem.
A test program, utilizing a large scale model, was run in the NASA Lewis Research Center 10- by 10-ft wind tunnel to examine the influence on performance of design parameters of turboprop S-duct inlet/diffuser systems. The parametric test program investigated inlet lip thickness, inlet/diffuser cross-sectional geometry, throat design Mach number, and shaft fairing shape. The test program was run at angles of attack to 15 deg and tunnel Mach numbers to 0.35. Results of the program indicate that current design techniques can be used to design inlet/diffuser systems with acceptable total pressure recovery, but several of the design parameters, notably lip thickness (contraction ratio) and shaft fairing cross section, must be optimized to prevent excessive distortion at the compressor face.
The analysis performed on the Space Shuttle Main Engine (SSME) High Pressure Fuel Turbopump (HPFTP) inlet housings is summarized. Three DIAL finite element models were build to aid in assessing the structural life of the welds and fillets at the vanes. Complete results are given.
The need for quantitative data is addressed by laser holographic interferometry (LHI) by efficiently recording and reducing high resolution density data. This paper investigates the potential to get quantitative data from an interferogram created from two separate holographic plates. THe LHI measurements were made on a scramjet inlet at Mach 4. LHI successfully generated results for both internal and external flows, and demonstrated its ability to map entire flow fields. Of the three data reduction techniques used, the four-bucket method gave better results than the three-bucket or three-by-three methods. This technique requires high quality holograms as well as a good resolution in the digital images to produce good data.
Calculations of the onset of flow instability are carried out for low-speed multistage axial compressors operating with asymmetric inlet flow. The modeling of the fluid dynamic interaction between the spoiled and unspoiled sectors of the compressor is the most important feature of the calculation procedure. The calculations show that annulus average slope of the compressor pressure rise characteristic equal to zero is a useful approximate stability criterion for situations where the dynamics of the compressor flow field do not couple strongly to the compression system or the structure of the imposed distortion is not similar to that of the eigenmodes of the flow in the compressor annulus. This criterion is employed to investigate the relationship between the present model and the 'parallel compressor' model. Calculations are also presented for cases when compressor and compressor system are closely coupled, and situations in which the compressor is subjected to a rotating distortion. These first-of-a-kind computations, and the accompanying description of the physical mechanisms, show that the stability of the flow in the compressor can be adversely affected if the temporal or spatial structure of the distortion is such that resonant type responses can be evoked either from the compressor or from compressor/compression system interactions.
Three-dimensional sidewall compression scramjet inlets with leading edge sweeps of 30 and 70 degrees have been tested in the Langley Hypersonic CF4 Tunnel at Mach 6 and a ratio of specific heats of 1.2. The effects of cowl position, contraction ratio, and Reynolds number were investigated. The models were instrumented with 42 static pressure orifices distributed on the sidewalls, baseplate, and cowl. Schlieren movies were made of each test for flow visualization of the entrance plane and cowl region. In order to obtain an approximate characterization of the flow field, a modification to two-dimensional inviscid oblique shock theory was derived to accommodate the three-dimensional effects of leading edge sweep. This theory qualitatively predicted the reflected shock structure/sidewall impingement locations and the observed increase in spillage (flow upturning) with increasing leading edge sweep. The primary effect of moving the cowl forward is capturing the flow which would have otherwise spilled out ahead of the cowl. Increasing the contraction ratio (moving the sidewalls closer together) increases the number of internal shock reflections and hence incrementally increases the sidewall pressure distribution. Significant Reynolds number effects were noted over a small range of Reynolds number.