Subsonic Blade Research Articles

A Coupled Effect Model of Two-Position Local Geometric Deviations on Subsonic Blade Aerodynamic Performance

The coupled aerodynamic performance influence of multiple deviations in the blade profile is a complex phenomena. In this paper, a high-pressure compressor rotor blade mid-span profile is studied with the design of experiments (DOE), numerical simulation and surrogate model to analyze the influence of the deviations on the blade aerodynamic performance. The purpose of this work is to provide a new rapid evaluation approach for the blade aerodynamic performance under multiple geometric deviations influence. First, the Hicks-Henne function was used to model the local geometric deviations of the blade profile, and the blade aerodynamic performance was calculated by using the computational fluid dynamics tools. By analyzing the calculation results, the momentum thickness of the boundary layer, the deviations height and the distance between the deviations are combined into a coupled effect model. Then, the coupled effect model was used to rapidly evaluate the blade aerodynamic performance when the two-position local geometric deviations exist on the blade surface. Finally, the evaluated performance were compared with the results predicted by a high-precision surrogate model, which verifies the high accuracy of the coupled effect model in evaluating the positive incidence range of the blade profile.

Full Annulus Simulations of a Transonic Axial Compressor Stage with Distorted Inflow at Transonic and Subsonic Blade Tip Speed

This article reports on systematic numerical studies on an axial compressor stage with distorted inflow. Four operating points at two speedlines were simulated with an inflow distortion generated by a 120 ∘ -sector segment with a wedge-type cross-section. With this setup, the interaction between the distorted inflow and the rotor flow was studied. The focus was put on the differences of the interaction between the distorted inflow and rotor flow as well as on the compressor behaviour at subsonic and transonic blade tip speeds, as the general mechanisms have been analysed in more detail in previous publications. The distorted flow itself is not influenced by the blade tip speed, but the interaction phenomena depend strongly on the spool speed and operating point. Additionally, the blade tip speed influences the circumferential sector of the compressor stage exit, which is affected by the distorted flow. The impact reaches from a small sector at 65% spool speed, with the peak efficiency operating point up to nearly the entire annulus at 100% spool speed, near the stall operating point.

Goal-oriented, model-constrained optimization for reduction of large-scale systems

Optimization-oriented reduced-order models should target a particular output functional, span an applicable range of dynamic and parametric inputs, and respect the underlying governing equations of the system. To achieve this goal, we present an approach for determining a projection basis that uses a goal-oriented, model-constrained optimization framework. The mathematical framework permits consideration of general dynamical systems with general parametric variations and is applicable to both linear and nonlinear systems. Results for a simple linear model problem of the two-dimensional heat equation demonstrate the ability of the goal-oriented approach to target a particular output functional of interest. Application of the methodology to a more challenging example of a subsonic blade row governed by the unsteady Euler flow equations shows a significant advantage of the new method over the proper orthogonal decomposition.

High-Resolution Large-Eddy Simulation of Flow Around Low-Pressure Turbine Blade

Large-eddy simulation ofcompressible Navier-Stokes equations is used to study flows where a laminar boundary-layer separation is followed by a turbulent reattachment. The aim of the present work is to predict and describe the transition process and its interaction with the wake dynamics for a subsonic blade turbine configuration. Indeed, a better knowledge of this mechanism can help to improve the accuracy of a Reynolds-averaged Navier-Stokes turbulence model for such a flow case. High-resolution large-eddy-simulation-type computations have been carried out for the T106 low-pressure blade turbine at inlet Mach number of 0.1 and chord Reynolds number of 1.6 x 10 5 based on the exit isentropic velocity. The simulated mean and turbulent quantities compare well with the available experimental data. The primary two-dimensional instability that originates from the free shear in the bubble is unstable via the Kelvin-Helmholtz mechanism. Then, the three-dimensional motions spread on the boundary layer, leading to full breakdown to turbulence after the reattachment point

Evolution of Upstream Propagating Shock Waves From a Transonic Compressor Rotor

The evolution of upstream propagating shock waves from the isolated transonic compressor designated NASA Rotor-35 is examined numerically. Results from the numerical simulations are compared with those from a semi-analytical two-dimensional model based on the nonlinear acoustic interaction of shock waves in the axial-tangential plane upstream of the rotor. The evolution determined from a two-dimensional viscous computational solution is found to agree well with the semi-analytical prediction and confirms that shock wave evolution is a primarily inviscid phenomenon. Radial variations are found to increase the rate of decay of the shock wave amplitude in comparison to the prediction from the semi-analytical two-dimensional model. The velocity field from the three-dimensional viscous solution compares well with experimental measurements, indicating that the initial shock strength and shock wave evolution immediately upstream of the rotor blade leading edge are accurately captured. The upstream-propagating shock system is found responsible for nearly 20% of the total loss attributable to the rotor, and is consistent with earlier transonic airfoil cascade studies. The axial decay rate of the upstream induced circumferential static pressure distortion is found to be an order of magnitude slower at spanwise locations with supersonic relative inlet Mach numbers than those at which it is subsonic. As a consequence of this slower decay rate, it is found that the axial gap to the upstream stator would need to be about twice that used for subsonic blade sections.

Active Control of Turbomachine Discrete Frequency Noise Utilizing Oscillating Flaps and Pistons

Turbomachine discrete frequency tones are a signi® cant environmental concern. These tones are generated by periodic blade row interactions, with only speci® c circumferential acoustic modes generated. This research is directed at active control of discrete frequency noise generated by subsonic blade rows through cancellation of the propagating acoustic waves, accomplished by utilizing airfoil surface mounted oscillating actuators to generate additional control propagating pressure waves. These control waves interact with the propagating acoustic waves, thereby canceling the acoustic waves and thus the far-® eld discrete frequency tones. A series of experiments are described, directed at investigating the fundamentals of this active discrete frequency noise control technique as well as verifying basic modeling assumptions. Both an isolated vane and a three-vane row are utilized, with the

An asymptotic theory of near-field propeller acoustics

This paper presents expressions for the harmonic components of the near-field acoustic pressure of a B-bladed unswept single-rotation propeller. These are derived using asymptotic approximations to the standard radiation integrals for steady loading and thickness noise, under the assumption that B is large. The dependence of the pressure on blade operating conditions (both supersonic and subsonic) is described by simple formulae, which provide significant insights into the mechanisms of sound generation by rotating bodies. For supersonic motion, the importance of sources satisfying the Ffowcs Williams & Hawkings sonic condition is demonstrated, whilst for subsonic blades the near-field noise is proved to be tip-dominated. Expressions for the noise (valid from close to the tips right out to infinity) are given in both cases, requiring matching across an Airy function smoothing region when the tips move subsonically. Excellent agreement between the asymptotic formulae and both full numerical evaluations (with a considerable saving in CPU time) and experimental data is achieved.

Design of Turbine Blades Suitable for Supersonic Relative Inlet Velocities and the Investigation of Their Performance in Cascades: Part II-Experiments, Results and Discussion

This paper describes cascade tests which were initially done to see whether blades designed by conventional methods for subsonic flows would in fact accept supersonic velocities at inlet to the blades but it was found that they would not do so. An analytical method for the design of turbine blades to accept supersonic relative inlet velocities has been presented in Part I (1). A blade section was designed for isentropic flow and from tests run on a cascade of these, values for losses and allowable contraction ratios were obtained which were incorporated in a revised design method. The tests and results obtained from the first supersonic cascade are described and also the results from the modified design which was based on non-isentropic conditions. The supersonic cascades were tested over a range of incidences in a small wind tunnel fitted with interchangeable liners to produce Mach numbers of approximately 1.5 and 1.3, Preliminary tests were also run at subsonic velocities. At design conditions flow-visualization by means of a Schlieren system showed wave formations in the blade passages which corresponded to the theoretical design values thus proving the validity of the design method. At off-design conditions a somewhat unusual shock-wave boundary-layer interaction was observed which appears to be associated with blades of this type. This was analysed by means of total head traverses, static pressure measurements and by the method of characteristics, and the effect which the configuration had on the downstream flow and losses is discussed. Total head traverses were done at the exits from the blade passages and also downstream of the trailing edges and the velocity coefficients were deduced from these values and from static pressure measurements obtained from wall tappings. For the first cascade which had a flow turning angle of 128° the velocity coefficient was approximately 0.95 and for the last cascade which had a turning angle some 9° greater, the value was 0.94. These velocity coefficients are higher than those for subsonic blades at comparative turning angles. With blades designed for supersonic flows by the method outlined in Part I (1) it appears to be possible to turn the flow through large angles such as 140° and to obtain acceptably high efficiencies. Tests in steam turbines indicate that with these blades the maximum efficiency is at least equal to that of a turbine designed for subsonic flow but that the heat drop is some 12 per cent higher.

Supersonic Cascade Tunnel Used to Evaluate Compressor Blade Performance

Large improvements in specific fuel consumption and weight have been achieved during the development of the aircraft gas turbine engine in the past decade. The turbo-machinery requirements have become more demanding during this period, resulting in a significant increase in relative Mach number and diffusion loading in fan and compressor stages. Considerable effort was devoted to develop design systems for this type of turbomachinery, in which relative supersonic inlet velocities are realized for rotor and stator blades. One of the major requirements was to develop a design system for the selection of supersonic blade elements which are required for fan and compressor stages employed in turbofan engines. Since subsonic blade selection design systems are based primarily on empirical relationships which are derived from two-dimensional cascade data, a two-dimensional cascade tunnel was developed for: (a) the investigation of supersonic cascade aerodynamics; (b) the development of high performance supersonic blade elements; and (c) the production of basic supersonic cascade data, which is required for the extrapolation of current subsonic design systems. The present paper describes this facility, presents some typical data, and discusses some of the results obtained.