Miklos Sajben

Forced oscillation experiments in supercritical diffuser flows

Low-frequency oscillations induced in ramjet inlets by combustion instabilities were simulated by mechanically modulating the exit area of a two-dimensional, supercritical diffuser at frequencies up to 330 Hz. Boundary layers were attached below a terminal shock Mach number of 1.27, and shock-induced separation occurred above this value up to the experimental limit of 1.35. Shock position histories were obtained and streamwise distributions of static/total dynamic pressures were determined both on the wall and within the flow for various shock strengths and frequencies. Excitation at the natural frequencies of the shock motion produced no obvious resonance effects. For weak shocks, the perturbations and their reflections from the shock are reasonably approximated by one-dimension al, acoustic considerations, but this description fails for strong shocks.

Experimental investigation of terminal shock sensors for mixed-compression inlets

This paper describes experimental investigations of devices designed for the nonintrusive detection of terminal shock location in mixed-compression inlets at high supersonic flight speeds. Systems based on sensing wall pressures by an array of wall-mounted transducers were selected for detailed study. Pressure signals were processed by three different methods: (1) interpretation of instantaneous pressure distributions, (2) detection of the turbulent intensity amplification occurring at the shock, and (3) determination of the upstream limit to which a search-tone, introduced at the downstream end of the channel, can propagate. The first two of these methods were tested in real time. The third method appeared feasible for weak shocks only; at high shock strengths, propagation upstream of the source could not be detected.

Response of a supersonic inlet to downstream perturbations

Experimental results are reported for flows in a ramp-type, external compression inlet with a large-aspectratio, rectangular cross section. The inlet was operated at a freestream Mach number of 1.84 with mechanically generated downstream perturbations. High-speed schlieren and time-dependent pressure measurements were employed extensively. In supercritical operation, pressure fluctuations throughout the inlet caused by the excitation varied linearly with the fluctuations at the exit station, even for large exit station amplitudes. In subcritical operation (buzz), the excitation interacted nonlinearly with the naturally present, highly periodic oscillations by either modifying the natural frequency, if the excitation was near a natural harmonic, or by having the excitation modulate the naturally occurring oscillation. In addition, the conditions at the two criticality boundaries were determined as a function of excitation amplitude and frequency.

Experimental investigation of normal-shock/turbulent-boundary-layer interactions with and without mass removal

Interactions of a normal shock with a turbulent boundary layer over a flat surface were investigated with and without imposing mass removal. The approach flow was very nearly two-dimensional, with a uniform freestream Mach number of 1.48, strong enough to cause separation at the shock foot. Suction was imposed immediately upstream of the shock over a finely perforated surface, removing much of the boundary-layer flow. The time-mean wall pressure distributions and two components of the time-mean velocity vector field were determined for both cases. Mass removal distorted the original shock pattern and eliminated the separation at the shock foot while reducing the boundary-layer thicknesses and growth rates to less than half. Mass removal also introduced some undesirable effects: asymmetric sidewall boundary-layer growth in the subsonic flow and increased shock oscillation amplitudes. The removed mass flow varied strongly in the streamwise direction in a manner that indicated wide variations of the flow coefficients for the individual perforations.

Experimental study of flows in a two-dimensional inlet model

Experimental results are reported for flows in a ramp-type, external-compression inlet with a rectangular, large-aspect-ratio cross section, at a freestream Mach number of 1.84. Variation of the exit throttle area created a wide range of operating conditions, from highly supercritical to fully subsonic internal flows with a detached shock ahead of the ramp lip. Both time-mean and dynamic aspects of the flows were investigated. At high terminal-shock Mach numbers (1.5-2.2), three different, massively separated, subsonic flow patterns were found, which depended on throttle setting and the history of prior flow conditions. A specific pressure ratio could be associated with at least two such patterns. At low pressure ratios, the terminal shock was weak and strongly influenced by the ramp/cowl configuration. The presence of leading edges is believed to have been closely involved in the large-amplitude, periodic oscillations (buzz) observed at the low end of the pressure-ratio range.

Reflection of Large Amplitude Acoustic Pulses from an Axial Flow Compressor

The reliability of unsteady inlet flow computations may be seriously degraded by the lack of experimentally validated compressor-face (outflow) boundary conditions. It is common practice to require some flow variable to be constant (pressure, velocity or Mach number, etc.), but there is little if any documented evidence to support these assumptions. In addition, these assumptions do not have a basis in observation nor are they likely to actually occur during a rapid transient of a real inlet/engine system. The present paper presents measurements of acoustic reflection coefficients for an operating multistage compressor, a quantity appropriate for the characterization of the compressor face for computational purposes. The experiment used an impulse method, in which shortduration, large amplitude acoustic pulses (one msec, with a peak value of nearly 4% of the mean static pressure, respectively) were generated in an constantarea, annular inlet. The pulse and its reflection from the compressor-face were tracked by fast-response pressure transducers. Frequency-domain analysis of the data yields transfer functions that may be thought of as frequency-resolved reflection coefficients. It was demonstrated that transfer functions can be used for the accurate prediction of transients induced by arbitrary acoustic incident disturbances. None of the currently imposed compressor-face boundary conditions predict the data obtained in this study, indicating that current practices concerning outflow boundary conditions are in need of revision. Nomenclature

Experiment to Support the Formulation and Validation of Compressor-Face Boundary Conditions

The reliability of unsteady inlet e ow computations may be seriously degraded by the lack of experimentally validated compressor-face (outeow) boundary conditions. The commonly imposed oute ow conditions require a e ow variableto beconstant (pressure, velocity, Mach number,etc. )at theoute owboundary, butthereis littleif any documented evidenceto support theseassumptions, norarethey likely to actually occurduring a rapid transient of arealinlet/enginesystem.Measurementsarepresentedofacousticree ectioncoefe cientsforanoperatingmultistage compressor, a quantity appropriate for the characterization of the compressor face for computational purposes. The experiment used an impulse method, in which short-duration, large-amplitude acoustic pulses (1 ms, with a peak value of nearly 4% of the mean static pressure, respectively ) were generated in a constant-area, annular inlet. The pulse and its ree ection from the compressor face were tracked by fast-response pressure transducers. Frequency-domain analysis of the data yields transfer functions that may be thought of as frequency-resolved ree ection coefe cients. None of the currently availablecompressor-face boundary conditions accurately predict the data obtained in this study, indicating that current practices concerning oute ow boundary conditions are in need of revision.

Confined Normal-Shock/Turbulent-Boundary-Layer Interaction Followed by an Adverse Pressure Gradient

A steady, nominally two-dimensional interaction of a normal shock with a turbulent boundary layer over a flat surface is investigated experimentally. The approach Mach number is 1.34 and the Reynolds number based on the momentum thickness of the approach boundary layer is 14,600. The boundary layer displays intermittent flow reversal (incipient detachment) at the foot of the shock. The experiment differs from similar past studies by the imposition of an extended adverse pressure gradient region downstream of the shock and by a relatively high ratio of approach boundary-layer thickness to channel height. The time-mean and fluctuating velocity fields were explored in detail using laser Doppler velocimetry. Spatial distributions of turbulence kinetic energy, shear stress, and turbulence production are presented.

Inlet/Compressor System Response to Short-Duration Acoustic Disturbances

Recent studies have shown that computations of unsteady high-speed inlet flows produce results that strongly depend on the compressor-face boundary condition. Traditionally applied boundary conditions have no experimentalverification, which leads to uncertainty and significant risks in development programs. Experimental information is offered that can serve as a basis for constructing realistic boundary conditions. Short-duration transients were investigated in a constant-area circular inlet attached to an operating high-speed, single-stage, axial flow compressor. The transients were initiated by the generation of short-duration acoustic pulses within the inlet, using an exploding wire method. Pulse duration was typically 2 ms, with a peak amplitude of 3% of the mean inlet static pressure measured at the compressor face. Fast-response wall-mounted pressure transducers were used to detect the incident, reflected, and transmitted pulses. Data were obtained for axial compressor-face Mach numbers from 0.15 to 0.45. Frequency-domain analysis was used to extract dimensionless transfer functions that may be viewed as frequency-resolved reflection and transmission coefficients. The information is appropriate for the characterization of the compressor face for computational purposes. None of the customary boundary conditions predict the data obtained in this study, highlighting the need to revise conventional methods of imposing outflow boundary conditions.

Prediction of Acoustic, Vorticity and Entropy Waves Generated by Short-Duration Acoustic Pulses Incident on a Blade Row

Some long-standing questions concerning the dynamic behavior of coupled Wet/compressor systems have been answered by analyzing data from recent experiments in which the reflections of intense, short-duration acoustic pulses from an operating compressor were documented. The present paper offers a simple, one-dimensional integral theory as a background for these experiments. The arrival of an acoustic pulse (or an acoustic step change) to a single row of stationary blades gives rise to two acoustic waves (one upstream and one downstream), one vorticity wave, and possibly also an entropy wave. Pulses are characterized by integrals of spatial distributions of pressure, temperature or tangential velocity, while steps are defined by the jump of these flow properties at the wave. Simple expressions are given for the strength of each wave type, in terms of the blade stagger angle and the axial Mach number. The results help in interpreting, scaling and extrapolating experimental data. The work is a step towards an ability to specify realistic outflow boundary conditions in unsteady inlet flow computations, as appropriate for the compressor geometry and the operating conditions. LIST OF SYMBOLS