An Inviscid-Viscous Interaction Method to Predict the Blade-to-Blade Performance of Axial Compressors

1980 ◽  
Vol 31 (3) ◽  
pp. 173-196 ◽  
Author(s):  
W.J. Calvert ◽  
M.V. Herbert
Keyword(s):  
Inviscid Flow ◽  
Surface Boundary ◽  
Interaction Method ◽  
Suction Surface ◽  
Surface Boundary Layer ◽  
Axial Compressors ◽  
Pressure Loss Coefficient ◽  
Viscous Interaction ◽  
Integral Methods ◽  
Time Marching

SummaryAn inviscid-viscous interaction method has been developed to predict the blade-to-blade flow in axial compressors. The method is primarily intended for high deflection, transonic (but substantially shock-free) blades where the suction surface boundary layer may be separated near the trailing edge. The inviscid flow is calculated by a time marching method and the viscous flow by integral methods for laminar and turbulent boundary layers. A mixing calculation is then carried out to determine the blade deviation angle and pressure loss coefficient. Predictions have been compared with test results for three high deflection, transonic cascades and there is generally good agreement.

An Inviscid-Viscous Interaction Treatment to Predict the Blade-to-Blade Performance of Axial Compressors With Leading Edge Normal Shock Waves

1982 ◽  
Author(s):  
W. J. Calvert
Keyword(s):  
Shock Wave ◽  
Leading Edge ◽  
Inviscid Flow ◽  
Normal Shock ◽  
Surface Boundary ◽  
Suction Surface ◽  
Axial Compressors ◽  
Viscous Interaction ◽  
Interaction Treatment ◽  
Inlet Conditions

An inviscid-viscous interaction treatment has been developed to predict the blade-to-blade flow in axial compressors operating with supersonic inlet conditions and a normal shock wave at inlet to the blade passage. The treatment uses both direct and inverse modes of operation for the inviscid and viscous calculations, and thus it can model the separation of the suction surface boundary layer produced by the strong interaction with the shock wave. The inviscid flow is calculated by a time marching method and the viscous flow by integral methods for laminar and turbulent boundary layers.

Boundary Layer Development in the BR710 and BR715 LP Turbines—The Implementation of High-Lift and Ultra-High-Lift Concepts

2002 ◽  
Vol 124 (3) ◽  
pp. 385-392 ◽  
Author(s):  
R. J. Howell ◽  
H. P. Hodson ◽  
V. Schulte ◽  
R. D. Stieger ◽  
Heinz-Peter Schiffer ◽  
...  
Keyword(s):  
Boundary Layer ◽  
Test Facility ◽  
Surface Boundary ◽  
Unsteady Boundary Layer ◽  
Suction Surface ◽  
High Lift ◽  
Surface Boundary Layer ◽  
Cold Flow ◽  
Layer Behavior ◽  
Lp Turbine

This paper describes a detailed study into the unsteady boundary layer behavior in two high-lift and one ultra-high-lift Rolls-Royce Deutschland LP turbines. The objectives of the paper are to show that high-lift and ultra-high-lift concepts have been successfully incorporated into the design of these new LP turbine profiles. Measurements from surface mounted hot film sensors were made in full size, cold flow test rigs at the altitude test facility at Stuttgart University. The LP turbine blade profiles are thought to be state of the art in terms of their lift and design philosophy. The two high-lift profiles represent slightly different styles of velocity distribution. The first high-lift profile comes from a two-stage LP turbine (the BR710 cold-flow, high-lift demonstrator rig). The second high-lift profile tested is from a three-stage machine (the BR715 LPT rig). The ultra-high-lift profile measurements come from a redesign of the BR715 LP turbine: this is designated the BR715UHL LP turbine. This ultra-high-lift profile represents a 12 percent reduction in blade numbers compared to the original BR715 turbine. The results from NGV2 on all of the turbines show “classical” unsteady boundary layer behavior. The measurements from NGV3 (of both the BR715 and BR715UHL turbines) are more complicated, but can still be broken down into classical regions of wake-induced transition, natural transition and calming. The wakes from both upstream rotors and NGVs interact in a complicated manner, affecting the suction surface boundary layer of NGV3. This has important implications for the prediction of the flows on blade rows in multistage environments.

Transonic Relief in Fans and Compressors

2020 ◽  
Author(s):  
Demetrios Lefas ◽  
Robert J. Miller
Keyword(s):  
Mach Number ◽  
Physical Mechanism ◽  
Strong Shock ◽  
Area Ratio ◽  
Surface Boundary ◽  
Suction Surface ◽  
Surface Boundary Layer ◽  
Will Self ◽  
Optimal Area ◽  
Transonic Fan

Abstract Every supersonic fan or compressor blade row has a streamtube, the ‘sonic streamtube’, which operates with a blade relative inlet Mach number of one. A key parameter in the design of the ‘sonic streamtube’ is the area ratio between the blade throat area and upstream passage area, Athroat/Ainlet. In this paper, it is shown that one unique value exists for this area ratio. If the area ratio differs, even slightly, from this unique value then the blade either chokes or has its suction surface boundary layer separated due to a strong shock. It is therefore surprising that in practice designers have relatively little problem designing blade sections with an inlet relative Mach number close to unity. This paper shows that this occurs due to a physical mechanism known as ‘transonic relief’. If a designer makes a mistake, and designs a blade with a ‘sonic streamtube’ which has the wrong area ratio, then ‘transonic relief’, will self-adjust the spanwise streamtube height automatically moving it towards the unique optimal area ratio, correcting for the designer’s error. Furthermore, as the blade incidence changes, the spanwise streamtube height self-adjusts, moving the area ratio towards its unique optimal value. Without ‘transonic relief’, supersonic and transonic fan and compressor design would be impossible. The paper develops a simple model which allows ‘transonic relief’ to be decoupled from other mechanisms, and to be systematically studied. The physical mechanism on which it is based is thus determined and its implications for blade design and manufacturing discussed.

Physical Explanations of Tip Leakage Flow Field in an Axial Compressor Rotor

1998 ◽  
Author(s):  
Masahiro Inoue ◽  
Masato Furukawa ◽  
Kazuhisa Saiki ◽  
Kazutoyo Yamada
Keyword(s):  
Boundary Layer ◽  
Vortex Core ◽  
Leakage Flow ◽  
Surface Boundary ◽  
Suction Surface ◽  
Surface Boundary Layer ◽  
Tip Leakage Vortex ◽  
Tip Leakage ◽  
Compressor Rotor ◽  
Vortex Lines

Structure of a tip leakage flow field in an axial compressor rotor has been investigated by detailed numerical simulations and appropriate post-processing. Physical explanations of the structure are made in terms of vortex-core identification, normalized helicity, vortex-lines, limiting streamlines, etc. The onset of the discrete tip leakage vortex is located on the suction surface at some distance from the leading edge. The vortex core with high vorticity is generated from a shear layer between the leakage jet flow and the main flow. The streamlines in the leakage flow are coiling around the vortex core. All the vortex-lines in the tip leakage vortex core link to ones in the suction surface boundary layer. The other vortex-lines in the suction surface boundary layer link to the vortex-lines in the pressure surface boundary layer and in the casing wall boundary layer. There are two mechanisms to reduce intensity of the tip leakage vortex: one is reduction of discharged vorticity caused by the linkage of vortex-lines between the suction surface and casing wall boundary layers, and another is diffusion of vorticity from the tip leakage vortex. Relative motion of the endwall has a substantial influence on the structure of the leakage flow field. In the case of a compressor rotor, it intensifies streamwise vorticity of the leakage vortex but reduces leakage flow loss.

Predicting the Profile Loss of High-Lift Low Pressure Turbines

2011 ◽  
Vol 134 (2) ◽  
Author(s):  
John D. Coull ◽  
Howard P. Hodson
Keyword(s):  
Boundary Layer ◽  
Velocity Distribution ◽  
Low Pressure ◽  
Design Tool ◽  
Surface Boundary ◽  
Preliminary Design ◽  
Suction Surface ◽  
Surface Boundary Layer ◽  
Blade Loading ◽  
Profile Loss

The overall efficiency of low pressure turbines is largely determined by the two-dimensional profile loss, which is dominated by the contribution of the suction surface boundary layer. This boundary layer typically features a laminar separation bubble and is subjected to an inherently unsteady disturbance environment. The complexity of the flow behavior makes it difficult to numerically predict the profile loss. To address this problem, an empirical method is proposed for predicting the boundary layer integral parameters at the suction surface trailing edge, allowing the profile loss to be estimated. Extensive measurements have been conducted on a flat plate simulation of the suction surface boundary layer. The disturbance environment of real machines was modeled using a moving bar wake generator and a turbulence grid. From this data set, empirically based methods have been formulated using physical principles for the prediction of the momentum thickness and shape factor at the suction surface trailing edge. The predictions of these methods may be used to estimate the profile loss of a given cascade, which achieves reasonable agreement with the available data. By parameterizing the shape of the suction surface velocity distribution, the method is recast as a preliminary design tool. Powerfully, this may be used to guide the selection of the key design parameters (such as the blade loading and velocity distribution shape) and enables a reasonable estimation of the unsteady profile loss to be made at a very early stage of design. To illustrate the capabilities of the preliminary design tool, different styles of velocity distribution are evaluated for fixed blade loading and flow angles. The predictions suggest that relatively “flat-top” designs will have the lowest profile loss but good performance can also be achieved with front-loaded “peaky” distributions. The latter designs are more likely to have acceptable incidence tolerance.

Development of Controlled Diffusion Airfoils for Multistage Compressor Application

1984 ◽  
Vol 106 (2) ◽  
pp. 271-278 ◽  
Author(s):  
D. E. Hobbs ◽  
H. D. Weingold
Keyword(s):  
High Efficiency ◽  
Boundary Layer Separation ◽  
Surface Boundary ◽  
Standard Series ◽  
Suction Surface ◽  
Surface Boundary Layer ◽  
Controlled Diffusion ◽  
Multistage Compressor ◽  
Mach Numbers ◽  
Inlet Conditions

A series of Controlled Diffusion Airfoils has been developed for multistage compressor application. These airfoils are designed analytically to be shock-free at transonic Mach numbers and to avoid suction surface boundary layer separation for a range of inlet conditions necessary for stable compressor operation. They have demonstrated, in cascade testing, higher critical Mach numbers, higher incidence range, and higher loading capability than standard series airfoils designed for equivalent aerodynamic requirements. These airfoils have been shown, in single and multistage rig testing, to provide high efficiency, high loading capability, and ease of stage matching, leading to reduced development costs and improved surge margin. The Controlled Diffusion Airfoil profile shapes tend to have thicker leading and trailing edges than their standard series counterparts, which lead to improved compressor durability.

scholarly journals Direct numerical simulations of transition in a compressor cascade: the influence of free-stream turbulence

2010 ◽  
Vol 665 ◽  
pp. 57-98 ◽  
Author(s):  
TAMER A. ZAKI ◽  
JAN G. WISSINK ◽  
WOLFGANG RODI ◽  
PAUL A. DURBIN
Keyword(s):  
Boundary Layer ◽  
Pressure Gradient ◽  
Free Stream ◽  
Transition To Turbulence ◽  
Surface Boundary ◽  
Bypass Transition ◽  
Suction Surface ◽  
Surface Boundary Layer ◽  
Laminar Separation ◽  
Free Stream Turbulence

The flow through a compressor passage without and with incoming free-stream grid turbulence is simulated. At moderate Reynolds number, laminar-to-turbulence transition can take place on both sides of the aerofoil, but proceeds in distinctly different manners. The direct numerical simulations (DNS) of this flow reveal the mechanics of breakdown to turbulence on both surfaces of the blade. The pressure surface boundary layer undergoes laminar separation in the absence of free-stream disturbances. When exposed to free-stream forcing, the boundary layer remains attached due to transition to turbulence upstream of the laminar separation point. Three types of breakdowns are observed; they combine characteristics of natural and bypass transition. In particular, instability waves, which trace back to discrete modes of the base flow, can be observed, but their development is not independent of the Klebanoff distortions that are caused by free-stream turbulent forcing. At a higher turbulence intensity, the transition mechanism shifts to a purely bypass scenario. Unlike the pressure side, the suction surface boundary layer separates independent of the free-stream condition, be it laminar or a moderate free-stream turbulence of intensityTu~ 3%. Upstream of the separation, the amplification of the Klebanoff distortions is suppressed in the favourable pressure gradient (FPG) region. This suppression is in agreement with simulations of constant pressure gradient boundary layers. FPG is normally stabilizing with respect to bypass transition to turbulence, but is, thereby, unfavourable with respect to separation. Downstream of the FPG section, a strong adverse pressure gradient (APG) on the suction surface of the blade causes the laminar boundary layer to separate. The separation surface is modulated in the instantaneous fields of the Klebanoff distortion inside the shear layer, which consists of forward and backward jet-like perturbations. Separation is followed by breakdown to turbulence and reattachment. As the free-stream turbulence intensity is increased,Tu~ 6.5%, transitional turbulent patches are initiated, and interact with the downstream separated flow, causing local attachment. The calming effect, or delayed re-establishment of the boundary layer separation, is observed in the wake of the turbulent events.

Rotating stratified flow past a steep-sided obstacle: incipient separation

1989 ◽  
Vol 206 ◽  
pp. 47-73 ◽  
Author(s):  
M. R. Foster
Keyword(s):  
Boundary Layer ◽  
Narrow Band ◽  
Inviscid Flow ◽  
Small Region ◽  
Ekman Layer ◽  
Surface Boundary ◽  
Surface Boundary Layer ◽  
Viscous Shear ◽  
Incipient Separation ◽  
Taylor Column

Many of the most interesting phenomena observed to occur in the flow of rotating and stratified fluids past obstacles, for example eddy shedding and wake unsteadiness, are due to separation of the boundary layer on the obstacle or its Taylor column. If the Rossby number of the flow lies between E½ and E (E is the Ekman number) and the Burger number is small, the structure of a viscous shear layer of width E⅙ on the circumscribing cylinder of an axisymmetric obstacle controls the inviscid flow. The surface boundary layer is not an Ekman layer, but a Prandtl layer, even at small Rossby numbers. As the slope of the obstacle at its base increases, the nature of the inviscid motion is altered substantially, in the rotation-dominated regime. We show that, for sufficiently large slopes, the flow develops a small region of non-uniqueness external to the column, simultaneously with the separation of the narrow band of fluid flowing round the base of the object.

scholarly journals Computation of Transonic Flows in and About Turbine Cascades With Viscous Effects

1984 ◽  
Author(s):  
U. K. Singh
Keyword(s):  
Inviscid Flow ◽  
Viscous Effects ◽  
Trailing Edge ◽  
Pressure Surface ◽  
Integral Methods ◽  
Interaction Treatment ◽  
Time Marching ◽  
Flow Through ◽  
Turbine Cascades ◽  
Good Agreement

An inviscid-viscous interaction treatment has been developed to predict the flow through transonic axial turbine blade cascades. The treatment includes a trailing-edge base pressure model. This model is based on treating the area between the points of flow separation on the blade surfaces at the trailing-edge and the point of downstream confluence of the suction and pressure surface flows as a region of constant pressure. A time marching technique is used to calculate the inviscid flow and viscous flow is calculated by integral methods for laminar and turbulent boundary layers. Good agreement with experimental data has been obtained.

scholarly journals The Effect of Nonequilibrium Boundary Layers on Compressor Performance

2018 ◽  
Vol 140 (10) ◽  
Author(s):  
Andrew P. S. Wheeler ◽  
Anthony M. J. Dickens ◽  
Robert J. Miller
Keyword(s):  
Boundary Layers ◽  
Compressor Blade ◽  
Surface Boundary ◽  
Order Model ◽  
Suction Surface ◽  
Surface Boundary Layer ◽  
Reduced Order ◽  
Compressor Performance ◽  
Profile Loss ◽  
Made In

The paper investigates the effect of nonequilibrium behavior of boundary layers on the profile loss of a compressor. The investigation is undertaken using both high fidelity simulations of a midheight section of a compressor blade and a reduced order model, MISES. The solutions are validated using experimental measurements made in the embedded stage of a multistage low speed compressor. The paper shows that up to 35% of the suction surface boundary layer of the compressor blade exhibits nonequilibrium behavior. The size of this region reduces as the Reynolds number is increased. The nonequilibrium behavior was found to reduce profile loss in cases of attached transition and raise loss where transition occurs through separation.

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