Transonic Axial Flow Compressor Research Articles

An Iterative Method for Blade Profile Loss Model Adaptation Using Streamline Curvature

The various incidence, deviation, and loss models used in through-flow analysis methods, such as streamline curvature, are nothing more than statistical curve fits. A closer look at public domain data reveals that these statistical correlations and curve fits are usually based on experimental cascade data that actually display a fairly large scatter, resulting in a relatively high degree of uncertainty. This usually leads to substantial differences between the calculated and actual performances of a given gas turbine engine component. Typically, matching calculated results from a through-flow analysis against experimental data requires the combination of various correlations available in the public domain, through a very tedious, complex, and time consuming “trial and error” process. This particular study supports the view that it might actually be much more time effective to “adopt” a given loss model against experimental data through an iterative, physics-based approach, rather than try to identify the best combination of available correlations. For example, the well-established “Swan’s model” for calculating the blade profile loss factor in subsonic and transonic axial flow compressors depends strongly on approximate correlations for calculating the blade wake momentum thickness, and therefore represents such a case. This study demonstrates this by looking into an iterative approach to blade profile loss model adaptation that can provide a relatively simple and quick, but also physics-based way of “calibrating” profile loss models against available experimental data for subsonic applications. This paper presents in detail all the analysis necessary to support the above concept and discusses Swan’s model in particular as an example. Finally, the paper discusses the performance comparison of a two-dimensional, streamline curvature compressor model against experimental data before and after the adaptation of that particular loss model. This analysis proves the potential of the simulation strategy presented in this paper to “adopt” a given loss model against experimental data through an iterative, physics-based approach.

Aerodynamics of swept and leaned transonic compressor-rotors

A systematic investigation to understand the impact of axially swept and tangentially leaned blades on the aerodynamic behaviour of transonic axial-flow compressor rotors was undertaken. Effects of axial and tangential blade curvature were analyzed separately. A commercial CFD package, which solves the Reynolds-averaged Navier–Stokes equations, was used to compute the complex flow field of transonic compressor-rotors. It was validated against NASA Rotor 37 existing experimental data. Computed performance maps and downstream profiles showed a good agreement with measured ones. Furthermore, comparisons with experimental data indicated that the overall features of three-dimensional shock structure, shock-boundary layer interaction, and wake development are calculated well by the numerical solution. Next, quite a large number of new transonic swept rotors (26) were modelled from the original Rotor 37, by changing the meridional curvature of the original stacking line through three previously defined control points (located at 33%, 67% and 100% of span). Similarly, 26 new transonic leaned rotors were modelled by changing the circumferential position of the same control points. All the new transonic rotors were simulated and the results revealed many interesting aspects which are believed to be very helpful to better understand the blade curvature effects on shock structure and secondary losses within a transonic rotor.

Recirculation casing treatment by using a vaned passage for a transonic axial-flow compressor

The tip leakage flow has an important influence on compressor performance particularly for the case of transonic compressors. Recirculation casing treatment is an effective approach to control the tip clearance flow which results in stability enhancement. Numerical simulations for a high-speed axial-flow compressor rotor with endwall recirculation, applying a state of the art design for the recirculation passageway are presented in the current study. Anti-swirl vanes are placed in the recirculation passageway. The effects of recirculation casing treatment on the compressor overall characteristics, stability, blade loading, and flow field are presented and discussed. Results show a significant range extension with a very small penalty in efficiency. Besides, the recirculated mass flow is only a few per cent of the main flowrate.

Parametric Study of Tip Clearance—Casing Treatment on Performance and Stability of a Transonic Axial Compressor

The control of tip leakage flow through the clearance gap between the moving and stationary components of rotating machines is still a high-leverage area for improvement of stability and performance of aircraft engines. Losses in the form of flow separation, stall, and reduced rotor work efficiency are results of the tip leakage vortex (TLV) generated by interaction of the main flow and the tip leakage jet induced by the blade pressure difference. The effects are more detrimental in transonic compressors due to the interaction of shock TLV. It has been previously shown that the use of slots and grooves in the casing over tip of the compressor blades, known as casing treatment, can substantially increase the stable flow range and therefore the safety of the system but generally with some efficiency penalties. This paper presents a numerical parametric study of tip clearance coupled with casing treatment for a transonic axial-flow compressor NASA Rotor 37. Compressor characteristics have been compared to the experimental results for smooth casing with a 0.356 mm tip clearance and show fairly good agreement. Casing treatments were found to be an effective means of reducing the negative effects of tip gap flow and vortex, resulting in improved performance and stability. The present work provides guidelines for improvement of steady-state performance of the transonic axial-flow compressors and improvement of the stable operating range of the system.

Numerical Calculation of Three-Dimensional F1ow through A Transonic Compressor Rotor

Three-dimensional flow analysis is implemented to investigate the flow through transonic axial-flow compressor rotor(NASA R67) and to evaluate the performances of Abid's low-Reynolds-number k- and Baldwin-Lomax turbulence models. A finite volume method is used fur spatial discretization. The equations are solved implicitly in time by the use of approximate factorization. The upwind difference scheme is used for inviscid terms and viscous terms are approximated with central difference. The flux-difference-splitting method of Roe is used to obtain fluxes at the cell faces. Numerical analysis is performed near peak efficiency and near stall. The results are compared with the experimental data for NASA R67 rotor. Blade-to-Blade Mach number distributions are compared to confirm the accuracy of the code. From the results, it is concluded that Abid'k- model is better for the calculation of flow rate and efficiency than Baldwin-Lomax model. But, the predictions for Mach number and shock structure are almost the same.

Rotating Stall Control in a High-Speed Stage With Inlet Distortion: Part II—Circumferential Distortion

This paper presents the first attempt to stabilize rotating stall in a single-stage transonic axial flow compressor with inlet distortion using active feedback control. The experiments were conducted at the NASA Lewis Research Center on a single-stage transonic core compressor inlet stage. An array of 12 jet injectors located upstream of the compressor was used for forced response testing and feedback stabilization. Results for a circumferential total pressure distortion of about one dynamic head and a 120 deg extent (DC(60) = 0.61) are reported in this paper. Part I (Spakovszky et al., 1999) reports results for radial distortion. Control laws were designed using empirical transfer function estimates determined from forced response results. Distortion introduces coupling between the harmonics of circumferential pressure perturbations, requiring multivariable identification and control design techniques. The compressor response displayed a strong first spatial harmonic, dominated by the well-known incompressible Moore–Greitzer mode. Steady axisymmetric injection of 4 percent of the compressor mass flow resulted in a 6.2 percent reduction of stalling mass flow. Constant gain feedback, using unsteady asymmetric injection, yielded a further range extension of 9 percent. A more sophisticated robust H∞ controller allowed a reduction in stalling mass flow of 10.2 percent relative to steady injection, yielding a total reduction in stalling mass flow of 16.4 percent.

Rotating Stall Control in a High-Speed Stage With Inlet Distortion: Part I—Radial Distortion

This paper presents the first attempt to stabilize rotating stall in a single-stage transonic axial flow compressor with inlet distortion using active feedback control. The experiments were conducted at the NASA Lewis Research Center on a single-stage transonic core compressor inlet stage. An annular array of 12 jet-injectors located upstream of the rotor tip was used for forced response testing and to extend the compressor stable operating range. Results for radial distortion are reported in this paper. First, the effects of radial distortion on the compressor performance and the dynamic behavior were investigated. Control laws were designed using empirical transfer function estimates determined from forced response results. The transfer functions indicated that the compressor dynamics are decoupled with radial inlet distortion, as they are for the case of undistorted inlet flow. Single-input-single-output (SISO) control strategies were therefore used for the radial distortion controller designs. Steady axisymmetric injection of 4 percent of the compressor mass flow resulted in a reduction in stalling mass flow of 9.7 percent relative to the case with inlet distortion and no injection. Use of a robust H∞ controller with unsteady nonaxisymmetric injection achieved a further reduction in stalling mass flow of 7.5 percent, resulting in a total reduction of 17.2 percent.

Experimental Investigation of Stepped Tip Gap Effects on the Performance of a Transonic Axial-Flow Compressor Rotor

The effects of stepped-tip gaps and clearance levels on the performance of a transonic axial-flow compressor rotor were experimentally determined. A two-stage compressor with no inlet guide vanes was tested in a modern transonic compressor research facility. The first-stage rotor was unswept and was tested for an optimum tip clearance with variations in stepped gaps machined into the casing near the aft tip region of the rotor. Nine causing geometries were investigated consisting of three step profiles at each of three clearance levels. For small and intermediate clearances, stepped tip gaps were found to improve pressure ratio, efficiency, and flow range for most operating conditions. At 100 percent design rotor speed, stepped tip gaps produced a doubling of mass flow range with as much as a 2.0 percent increase in mass flow and a 1.5 percent improvement in efficiency. This study provides guidelines for engineers to improve compressor performance for an existing design by applying an optimum casing profile.

Improved Blade Profile Loss and Deviation Angle Models for Advanced Transonic Compressor Bladings: Part II—A Model for Supersonic Flow

New blading concepts as used in modern transonic axial-flow compressors require improved loss and deviation angle correlations. The new model presented in this paper incorporates several elements and treats blade-row flows having subsonic and supersonic inlet conditions separately. The second part of the present report focuses on the extension of a well-known correlation for cascade losses at supersonic inlet flows. It was originally established for DCA bladings and is now modified to reflect the flow situation in blade rows having low-cambered, arbitrarily designed blades including precompression blades. Finally, the steady loss increase from subsonic to supersonic inlet-flow velocities demonstrates the matched performance of the different correlations of the new model.

Improved Blade Profile Loss and Deviation Angle Models for Advanced Transonic Compressor Bladings: Part I—A Model for Subsonic Flow

New blading concepts as used in modern transonic axial-flow compressors require improved loss and deviation angle correlations. The new model presented in this paper incorporates several elements and treats blade-row flows having subsonic and supersonic inlet conditions separately. In the first part of this paper two proved and well-established profile loss correlations for subsonic flows are extended to quasi-two-dimensional conditions and to custom-tailored blade designs. Instead of a deviation angle correlation, a simple method based on singularities is utilized. The comparison between the new model and a recently published model demonstrates the improved accuracy in prediction of cascade performance achieved by the new model.

Low Aspect Ratio Transonic Rotors: Part 1—Baseline Design and Performance

The analytical design and experimental test of a single-stage transonic axial-flow compressor are described. This design is the baseline of a compressor design study in which several blade design parameters have been systematically varied to determine their independent effects on compressor performance. The baseline design consisted of ruggedizing an existing compressor design that demonstrated outstanding aerodynamic performance, to correct some undesirable aeromechanical characteristics. The design study was performed by varying only one design parameter at a time, keeping the other design variables as close as possible to the baseline design. Specific design parameters of interest were those for which very few data were available to determine their sensitivity on compressor performance. This paper describes the baseline compressor design and its experimental performance. A detailed definition and flow analysis of the baseline design test point (used as the basis for all subsequent design variations) are provided.

A Navier-Stokes analysis of 3-D transonic cascade flow and its visualization.

The purpose of this paper is, first, to analyze 3-D transonic cascade flow by using the TVD Navier-Stokes code developed by the authors. Since the momentum equations of contravariant velocities are used in the code, the treatments of boundary conditions, especially the solid wall and the periodic boundaries for the 3-D cascade flow, become very easy. A Chakravarthy-Osher type TVD scheme modified by the authors is adopted in order to clearly simulate 3-D complex flow having shock waves and separation bubbles. The second purpose is to visualize the computed results by using the 3-D color graphic display. As numerical examples, the 3-D viscous flows thorough a transonic axial-flow compressor rotor with or without a tip clearance are computed, and the colored oli-flow patterns, 3-D particle paths, 3-D time lines which show the shock/boundary layer interactions and a leakage vortex near the casing wall are obtained. In particular, the mechanism of occurrence of the leakage vortex is investigated in detail.

Computation of Potential Flow on S2 Stream Surface for a Transonic Axial-Flow Compressor Rotor

A set of conservative full potential function equations governing the fluid flow along a given S2 streamsurface in a transonic axial compressor rotor was obtained. By the use of artificial density and a potential function/density iteration, this set of equations can be solved, and the passage shock on the S2 streamsurface can be captured. A computer program for this analysis problem has been developed and used to compute the flow field along a mean S2 streamsurface in the DFVLR transonic axial compressor rotor. A comparison of computed results with DFVLR L2F measurement at 100 percent design speed shows fairly good agreement.

A Computation and Comparison With Measurements of Transonic Flow in an Axial Compressor Stage With Shock and Boundary Layer Interaction

The flow field within a transonic axial flow compressor stage has been computed using a three-dimensional time-marching technique. Limited viscous effects are considered by including a calculation of the blade surface boundary layers. The boundary layer calculation forms an integral part of the whole computation scheme, which consists of, respectively: (i) inviscid Mach number calculations, (ii) blade surface boundary layer displacement thickness calculations, (iii) inviscid Mach number calculations with mass flow adjustment (based on the calculated displacement thicknesses) on the blade surfaces. The boundary layer computation is done by using integral calculation methods and has specifically been developed to account for a shock and boundary layer interaction (should one exist). Comparisons are made with measured results obtained with an advanced laser velocimeter. The calculated Mach number contours are in extremely good agreement with the experimental results. It is concluded that the calculation technique is a useful tool in the design of transonic axial flow turbomachines.

Efficient Laser Anemometer for Intra-Rotor Flow Mapping in Turbomachinery

Innovative features of the anemometer include: (1) a rapid and efficient data acquisition process, (2) a detailed real-time graphic display of the data being accumulated, and (3) input laser beam positioning that maximizes the size of the intra-rotor region being mapped. Results are presented that demonstrate the anemometer’s capability in flow mapping within a transonic axial-flow compressor rotor. Typically, a velocity profile, derived from 30,000 measurements along 1000 sequential circumferential positions covering 20 blade passages, can be obtained in 30 s. The use of fluorescent seed particles allows flow measurements near the rotor hub and the casing window.

Use of the Characteristic Method for the Prediction of the Three-Dimensional Flow Field in High Transonic Compressors

The strong pressure gradients that take place in high performance transonic axial flow compressors induce important viscous interaction effects. In order to estimate these effects and to calculate the flow field in the inviscid core, the perfect gas approach is necessary. A three-dimensional stationary characteristic method is used for such an approach, and numerical results are given for the flow field in: • the blade channel of a supersonic cascade with converging side walls, • the blade channel of an annular supersonic cascade, • the inlet region of a linear supersonic blade cascade with converging side walls.

Tracer Particle Flow in a Compressor Rotor Passage with Application to LDV

Theme T ASER-Doppler velocimetry (LDV) is a noncontact JLtoptical technique for determining the velocity of fluids by measuring the velocity of tracer particles entrained in the fluid rather than by measuring the velocity of the fluid itself. To help identify application guidelines of the LDV technique for the aircraft turbomachinery field, a theoretical analysis was conducted of the dynamic behavior of micron-size particles moving in the three-dimensional flowfield of a rotating transonic axial-flow air compressor rotor. The particle-to-gas velocity ratio and angular deviation relative to the gas were determined as functions of particle diameter, mass density, and radial position. It was found that the particles move essentially on gas stream surfaces and that their radial position within the blade channel is independent of their size and mass density. Velocity and angular deviation generally less than 1 % and 1 ° were achieved with 1 g/cc mass density tracer particles with diameters of 1/x or less. Contents The flow system analyzed is the three-dimensional gasparticle flow through a transonic axial-flow air compressor rotor. The rotor was designed for a weight flow per unit annulus area of 40.6 lbm/sec-ft2, a tip speed of 1393 fps, a total pressure and temperature ratio of 1.601 and 1.162 respectively, and a tip solidity of 1.30. The rotor has 44 blades, a leading-edge tip diameter of 1.65 ft, a hub-to-tip radius ratio of 0.504, and a rotational speed of 16,100 rpm. Gas-particle flow was analyzed in a blade channel solution region which is bounded by adjacent blade surfaces and by upstream and downstream boundaries along which the flow is assumed to be uniform. The region extends from a radius of 60 to 94.3% of the blade span as measured from the rotor hub. The flowfield was analyzed by first numerically1 determining the magnitude and direction of the gas velocity at all points of the blade channel solution region. Once the gas flowfield was established, the path followed by a particle and its velocity along that path were determined by numerically integrating the equations which govern the relative motion of a particle moving through the blade channel. The force required to accelerate the particle results from the viscous drag and is expressed in terms of an empirically determined drag coefficient2 which corrects the Stokes drag law for rarefaction, compressibility and inertial effects. It was assumed that the particles are noninteracting, uniformly distributed, and in sufficiently small concentration that they do not alter the gas properties.

Experimental Investigation of a Two Stage Transonic Axial-Flow Compressor

Experimental Investigation of a Two Stage Transonic Axial-Flow Compressor

An Experiment With Aspect Ratio as a Means of Extending the Useful Range of a Transonic Inlet Stage of an Axial Flow Compressor

An experiment in unstalled range of a transonic axial flow compressor stage is discussed. Two rotors of identical design, but of differing aspect ratio, are compared. The study suggests that some criteria other than blade chord Reynolds number be used to define chord lengths of transonic stages.

Some Investigations Into Transonic Axial-Flow Compressors With High Stage Load Coefficients and Low Degrees of Reaction

A design for a transonic axial-flow compressor is described, in which a rotor wheel with a low degree of reaction operates in subsonic flow and the stator operates in supersonic flow. A compressor was run as a single-stage compressor and an outlet pressure ratio equal to that of an orthodox compressor with 4 stages. Some discussion of another compressor in which both the rotor wheel and the stator operate in supersonic flow is also given.