R. P. Shreeve

Viscous Flow in a Controlled Diffusion Compressor Cascade With Increasing Incidence

A detailed two-component laser Doppler velocimeter mapping of the flow through a controlled diffusion compressor cascade at low Mach number (∼ 0.25) and Reynolds number of about 7×10 5 , at three inlet air angles from design to near stall, is reported. Results for pressure-side boundary layer and near-wake also are summarized. Information sufficient to allow preliminary assessment of viscous codes is tabulated

Comparison of calculated and experimental cascade performance for controlled-diffusion compressor stator blading

The mid-span section of a previously reported controlled-diffusion compressor stator has been experimentally evaluated in cascade. Measurements are taken over a range of incidence angles for blade chord Reynolds numbers from 470,000 to 690,000. Blade chord length is 12.7 cm, aspect ratio is 2.0, and solidity is 1.67. Measurements include conventional cascade performance parameters as well as blade surface pressures. Computations are made for the inviscid flow field, surface boundary layers, and loss for several of the blade inlet angle conditions, are compared against corresponding data.

Wake measurements and loss evaluation in a controlled diffusion compressor cascade

The results of two component laser-Doppler velocimeter (LDV) surveys made in the near wake (to one fifth chord) of a controlled diffusion (CD) compressor blade in a large-scale cascade wind tunnel are reported. The measurements were made at three positive incidence angles from near design to angles thought to approach stall. Comparisons were made with calibrated pressure probe and hot-wire wake measurements and good agreement was found. The flow was found to be fully attached at the trailing edge at all incidence angles and the wake profiles were found to be highly skewed. Despite the precision obtained in the wake velocity profiles, the blade loss could not be evaluated accurately without measurements of the pressure field. The blade trailing edge surface pressures and velocity profiles were found to be consistent with downstream pressure probe measurements of loss, allowing conclusions to be drawn concerning the design of the trailing edge.

Stator-Averaged, Rotor Blade-to-Blade Near-Wall Flow in a Multistage Axial Compressor With Tip Clearance Variation

This paper describes the effect of tip clearance changes on the pressure at the case wall of a second-stage rotor. Wall shear distributions under the rotor tip are also presented. The results show low-pressure areas extending along the rotor suction side but lying away from the blade. Pressure contours indicate the tangential loading at the tip is lower than predicted by two-dimensional calculations; however, the predicted loading is observed between the lowest pressures path in the passage and the blade pressure side

Case wall pressures in a multistage axial compressor with tip clearance variation

Large cyclical pressure excursions from stator-averaged values are observed on the suction side of a rotor blade at the case wall in a multistage compressor as the rotor moves relative to the stator. The pressure changes correlate well with stator relative position, occur in the passage away from the blade, and are only slightly modified by tip clearance changes. Close to the entry and exit of the tip gap, however, the pressures remain stable. The pressures near the gap are only slightly affected by stator proximity and primarily vary with tip gap height.

Investigation of a Supersonic Surface Jet—Hypersonic Flow Interaction with Axial Symmetry

An experimental investigation of axially symmetric slot injection of a supersonic jet into a hypersonic stream has been carried out A cone of 5° half angle was constructed in such a way that jet injection at six different angles could be provided. The throat areas and exit areas of the slots were the same in all cases, and provided a nominal jet Mach number of (2)*. The freestream Mach number was 6 in all cases. The boundary layer at the point of injection was fully turbulent Optical investigations were made, and surface pressure distributions were obtained. Stagnation temperature and stagnation pressure surveys were made downstream of the injection point, the results being reduced by digital computer to provide mass flux profiles. The results indicate that three basic types of interaction occur depending on the angle of the jet. Equations describing the displacement of the dividing streamline for each of the three interaction types are established empirically, and the streamline displacement is compared to separation and shock displacement scales. Jet mass entrainment is also examined. It is tentatively concluded that jets of intermediate angle hold promise of advanced propulsion applications, while jets at small angles provide stable films, as necessary in thermal protection applications.

A Transonic Compressor Stage: Part 1 — Experimental Results

The results of an experimental program to test a transonic compressor stage are presented. The stage was designed using computational fluid dynamic (CFD) techniques to minimize the use of empirical methods. The main performance characteristics of the stage are presented along with a description of the test rig. Some of the technical difficulties and solutions implemented in sustained testing of a transonic stage are outlined. Equations used to mass average the pressure and temperature ratio, to derive the isentropic efficiency and a method of measuring losses due to the case-wall are presented. Results from the experimental program for 70, 80, 90 and 100% speeds for rig open throttle to near stall are presented. Stall occurred at all speeds except 100%, where it was avoided due to rotor stress concerns. Quantitative performance maps of stage total-to-total pressure ratio and isentropic efficiency were determined. In addition, a map showing the losses due to the case-wall is presented. The total-to-total stagnation pressure and temperature profiles at the exit to the stage over the entire test range were measured and used to calculate the isentropic efficiencies. Static pressure measurements along the case-wall and hub-wall through the stage are presented. The experimental program has yielded data from a transonic compressor stage that can be used in the validation of CFD [1] codes and evaluation of the original design methods.© 2004 ASME

A Viscous Axisymmetric Throughflow Prediction Method for Multi-Stage Compressors

This paper describes a computer code which solves viscous axisymmetric flow through multistage compressors. The code incorporates the modelling of 3-D effects which result from secondary flow and mixing and lead to property changes in the streamwise and spanwise directions. The method requires no extra data for loss, deviation and blockage. The necessary input data are the geometry, upstream stagnation conditions, rotational speed and mass flow rate. Blade wakes and their decay are modelled. The secondary flow component of the mixing coefficient modifies the uniform part and the result is used in the turbulent diffusion terms of the equation of motion.The P&W 3S1 low aspect ratio 3 stage compressor and UTRC 2 stage research compressor are used for validation. Considering the complexity of the flow in the multi-stage environment, it was concluded that the method gives encouraging results at a very economical rate.Copyright

Velocity Vector Determination from Multiple-Sensor Pneumatic Probe Measurements

Theme D data for compressors are often derived from measurements made with multiple-sensor pneumatic probes. Although it is generally admitted that the presence of a probe alters the flow to some extent, few attempts have been made to evaluate and to correct probe readings for the interference that they produce. The goal of the present work was to measure the velocity distribution between blade rows of a small transonic compressor (11-in. o.d. and 2-in. blade height), operating subsonically. Different probes were used in the compressor annulus and were also applied to measure uniform temperature flows in pipes, free jets, and annular ducts. A useful method for representing probe characteristics and correcting for the presence of flow boundaries was devised. The probe calibration away from boundaries was represented by two interrelated polynomials for the pitch angle and velocity magnitude. In application, the probepressure measurements were reduced, using a computer offline and distribution of temperature (measured separately) to velocity magnitude and pitch angle. The departure of probe measurements near a flow boundary was then treated as being a self-induced flow effect. The effect was examined in known flows and represented analytically. The analytic representation of the calibration and of the corrections allowed a routine reduction of probe data from compressor surveys. The method can be applied generally to probes having four or five sensors for the measurement of pitch angle, yaw angle, and velocity magnitude. The measurements made with different probes near to flow boundaries showed that errors could be surprisingly large or very small depending on particular design features. A new probe which also incorporated a temperature sensor was designed and used successfully to transonic Mach numbers.

Experimental Investigation During Stall and Surge in a Transonic Fan Stage and Rotor-Only Configuration

High-speed pressure measurements of a transonic compressor rotor-stator stage and rotor-only configuration during stall and surge are presented. Rotational speed data showed the difference between the rotor-only case and rotor-stator stage. The rotor-only case stalled and remained stalled until the control throttle was opened. In the rotor-stator stage the compressor surged entering a cyclical stalling and then un-stalling pattern. An array of pressure probes was mounted in the case wall over the rotor for both configurations of the machine. The fast response probes were sampled at 196 608 Hz as the rotor was driven into stall. Inspection of the raw data signal allowed the size and speed of the stall cell during its growth to be investigated. Post-processing of the simultaneous signals of the casing pressure showed the development of the stall cell from the point of inception and allowed the structure of the stall cell to be viewed.Copyright