Turbine blade-tip clearance excitation forces

The results of an effort to assess the existing knowledge and plan the required experimentation in the area of turbine blade tip excitation forces is summarized. The work was carried out in three phases. The first was a literature search and evaluation, which served to highlight the state of the art and to expose the need for an articulated theoretical experimental effort to provide not only design data, but also a rational framework for their extrapolation to new configurations and regimes. The second phase was a start in this direction, in which several of the explicit or implicit assumptions contained in the usual formulations of the Alford force effect were removed and a rigorous linearized flow analysis of the behavior of a nonsymmetric actuator disc was carried out. In the third phase a preliminary design of a turbine test facility that would be used to measure both the excitation forces themselves and the flow patterns responsible for them were conducted and do so over a realistic range of dimensionless parameters

A theory of post-stall transients in multistage axial compression systems

A theory is presented for post stall transients in multistage axial compressors. The theory leads to a set of coupled first-order ordinary differential equations capable of describing the growth and possible decay of a rotating-stall cell during a compressor mass-flow transient. These changing flow features are shown to have a significant effect on the instantaneous compressor pumping characteristic during unsteady operation, and henace on the overall system behavior. It is also found from the theory that the ultimate mode of system response, stable rotating stall or surge, depends not only on the B parameter but also on other parameters, such as the compressor length-to-diameter ratio. Small values of this latter quantity tend to favor the occurrence of surge, as do large values of B. A limited parametric study is carried out to show the impact of the different system features on transient behavior. Based on analytical and numerical results, several specific topics are suggested for future research on post-stall transients

Helicopter rotor noise due to ingestion of atmospheric turbulence

A theoretical study was conducted to develop an analytical prediction method for helicopter main rotor noise due to the ingestion of atmospheric turbulence. This study incorporates an atmospheric turbulence model, a rotor mean flow contraction model and a rapid distortion turbulence model which together determine the statistics of the non-isotropic turbulence at the rotor plane. Inputs to the combined mean inflow and turbulence models are controlled by atmospheric wind characteristics and helicopter operating conditions. A generalized acoustic source model was used to predict the far field noise generated by the non-isotropic flow incident on the rotor. Absolute levels for acoustic spectra and directivity patterns were calculated for full scale helicopters, without the use of empirical or adjustable constants. Comparisons between isotropic and non-isotropic turbulence at the rotor face demonstrated pronounced differences in acoustic spectra. Turning and contraction of the flow for hover and low speed vertical ascent cases result in a 3 dB increase in the acoustic spectrum energy and a 10 dB increase in tone levels. Compared to trailing edge noise, turbulence ingestion noise is the dominant noise mechanism below approximately 30 rotor harmonics, while above 100 harmonics, trailing edge noise levels exceed turbulence ingestion noise by 25 dB

Stalled flow performance of a single stage transonic compressor

Originally presented as the first author's thesis, M.S. in the M.I.T. Dept. of Mechanical Engineering, 1978Includes bibliographical references (page 48)The stalled flow performance of a single stage transonic compressor is examined, using data from the MIT Gas Turbine Laboratory's Blowdown Compressor Facility. Measurements of the blowdown corrected weightflow are included, as well as stage exit static to inlet total pressure rise, and rotating stall cell measurements. A comparison of the flow blockage, as represented by stall cell circumferential extent, with the blowdown corrected weightflow is made. The 100% design speed stalled flow performance characteristic is presented. Radial traverse.data is also presented including flow angles, static and total pressures, and Mach number components.Supported by the NASA Lewis Research Center grant NGL 22-009-38

A Change of Variables to the Dual and Factorization of Composite Anomalous Jacobians

Changes of variables giving the dual model are constructed explicitly for sigma-models without isotropy. In particular, the jacobian is calculated to give the known results. The global aspects of the abelian case as well as some of those of the cases where the isometry group is simply connected are considered. Considering the anomalous case, we infer by a consistency argument that the `multiplicative anomaly' should be replaceable by adequate rules for factorization of composite jacobians. These rules are then generalized in a simple way for composite jacobians defined in spaces of different types. Implimentation of these rules then gives specific formulas for the anomally for semisimple algebras and also for solvable ones.Comment: 15 pages, no figures, Latex file, A treatment of the global aspects of the abelian and of semisimple duality groups are added. General formulas for the mixed anomaly are derive

Blade Loading Effects on Axial Turbine Tip Leakage Vortex Dynamics and Loss

umerical simulations have been carried out to define the loss generation mechanisms associated with tip leakage in un-shrouded axial turbines. Tip clearance vortex dynamics are a dominant feature of two mechanisms important in determining this loss: (i) decreased swirl velocity due to vortex line contraction in regions of decreasing axial velocity, i.e., adverse pressure gradient and (ii) vortex breakdown and reverse flow in the vortex core. The mixing losses behave differently from the conventional view of flow exiting a turbine tip clearance. More specifically, it is shown, through both control volume arguments and computations, that as a swirling leakage flow passes through a pressure rise, such as in the aft portion of the suction side of a turbine blade, the mixed-out loss can either decrease or increase. For turbines the latter typically occurs if the deceleration is large enough to initiate vortex breakdown, and it is demonstrated that this is the case in modern turbines. The effect of blade pressure distribution on clearance losses is illustrated through computational examination of two turbine blades, one with forward loading at the tip and one with aft loading. A 15% difference in leakage loss is found between the two, due to lower clearance vortex deceleration (lower core static pressure rise) with forward loading, and hence lower vortex breakdown loss. Additional computational experiments, carried out to define the effects of blade loading, incidence, and solidity, are found to be consistent with the proposed ideas linking blade pressure distribution, vortex breakdown and turbine tip leakage loss

Power Balance in Aerodynamic Flows

A control volume analysis of the compressible viscous flow about an aircraft is performed,including integrated propulsors and flow control systems. In contrast to most past analyses which have focused on forces and momentum flow, in particular thrust and drag, the present analysis focuses on mechanical power and kinetic energy flow. The result is a clear identification and quantification of all the power sources, power sinks, and their interactions which are present in any aerodynamic flow. The formulation does not require any separate definitions of thrust and drag, and hence it is especially useful for analysis and optimization of aerodynamic configurations which have tightly integrated propulsion and boundary layer control systems

Effects of non-axisymmetric tip clearance on axial compressor performance and stability

September 1997Statement of responsibility on title-page reads: M.B. Graf, T.S. Wong, E.M. Greitzer, F.E. Marble, C.S. Tan, H-W Shin, D.C. WislerIncludes bibliographical references (pages 34-35)The effects of circumferentially non-uniform tip clearance on axial compressor performance and stability have been investigated experimentally and analytically. A theoretical model for compressor behavior with non-axisymmetric tip clearance has been developed and used to design a series of first-of-a-kind experiments on a four-stage, low speed compressor. The experiments and computational results together show clearly the central physical features and controlling parameters of compressor response to non-axisymmetric tip clearance. It was found that the loss in stall margin was more severe than that estimated based on average clearance. The stall point was, in fact, closer to that obtained with uniform clearance at the maximum clearance level. The circumferential length scale of the tip clearance (and accompanying flow asymmetry) was an important factor in determining the stall margin reduction.For the same average clearance, the loss in peak pressure rise was 50% higher for an asymmetry with fundamental wavelength equal to the compressor circumference than with wavelength equal to one-half the circumference. The clearance asymmetry had much less of an effect on peak efficiency; the measured maximum efficiency decrease obtained was less than 0.4 percent compared to the 8% decrease in peak pressure rise due to the asymmetric clearance. The efficiency penalty due to non-axisymmetric tip clearance was thus close to that obtained with a uniform clearance at the circumferentially-averaged level. The theoretical model accurately captured the decreases in both steady-state pressure rise and stable operating range which are associated with clearance asymmetry.It also gave a good description of the observed trends of (i) increasing velocity asymmetry with decreasing compressor flow, and (ii) decreasing effect of clearance asymmetry with decreasing dominant wavelength of the clearance distribution. The time resolved data showed that the spatial structure of the pre-stall propagating disturbances in the compressor annulus was well represented and that the stability limiting process could be linked to the unsteady structure of these disturbance modes. The model was also utilized for parametric studies to define how compressor performance and stability is affected by the circumferential distribution of clearance, steady-state compressor pressure-rise characteristic, and system dynamic parameters. Sensitivity to clearance asymmetry was found to fall off strongly with the (asymmetry-related) reduced frequency and to increase with peak pressure rise and increasing curvature of the characteristic near the peak.Sponsored by the Air Force Office of Scientific Research, and the Air Force Aero Propulsion Technology (AFRAPT) Progra

Experimental and Computational Investigation for In-Line Boundary Layer Ingestion

The aerodynamic characteristics of an aft-body, in-line mounted, boundary layer ingesting, electric ducted fan, propulsion installation system has been investigated through experimental and computational analysis. A modular wind-tunnel model allows variation in the geometry of the propulsion installation system to be assessed, in combination with fan speed. Various experimental measurement techniques, including LDA, seven-hole-probe and surface pressures are employed. The propulsion installation system has also been investigated using RANS CFD and comparison with experimental data is presented. An investigation of the boundary conditions for efficiently representing the fan in CFD is described. Initial results show reasonably good agreement between CFD and experiment, in terms of velocity profiles and surface pressures, but highlight remaining differences for cases exhibiting flow separation

Effect of end-wall riblets on radial turbine performance

This paper presents a detailed study of the impact of manufacturing residual riblets at the rotor hub surface of a radial inflow turbine on the flow within the rotor passages and their contribution to drag reduction. Numerical analysis has been used to study the effects of those features at design point conditions. Riblets with different height and spacing have been examined to determine the riblet geometry where the maximum drag reduction is achieved. The relative height of the riblets to rotor inlet blade height was introduced to generalise the results. At the end of this study the results were compared with the available data in literature. It was found that the introduction of riblets could reduce the wall shear stress at the hub surface, while they contribute to increasing the streamwise vorticity within the rotor passage. For the geometries tested, the minimum drag was achieved using riblets with relative height hrel = 2.5% equivalent to 19.3 wall units. The results revealed that the spacing between riblets have a minor effect on their performance, this is due to the size of the streamwise vortex above the hub surface which will be discussed in this work