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.

scholarly journals The Effect of Non-Equilibrium Boundary Layers on Compressor Performance

2017 ◽  
Author(s):  
Andrew P. S. Wheeler ◽  
Anthony M. J. Dickens ◽  
Robert J. Miller
Keyword(s):  
Boundary Layers ◽  
Compressor Blade ◽  
Surface Boundary ◽  
Suction Surface ◽  
Surface Boundary Layer ◽  
Reduced Order ◽  
Compressor Performance ◽  
Profile Loss ◽  
Made In ◽  
Non Equilibrium

The paper investigates the effect of non-equilibrium behaviour of boundary layers on the profile loss of a compressor. The investigation is undertaken using both direct numerical simulation (DNS) of a mid-height 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 non-equilibrium behaviour. The size of this region reduces as the Reynolds number is increased. The non-equilibrium behaviour was found to reduce profile loss in most cases, however, in a range of cases where transition occurs through a small separation the presence of non-equilibrium behaviour was found to increase profile 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.

Predicting the Profile Loss of High-Lift Low Pressure Turbines

2010 ◽  
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 (LP) 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 dataset, 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. 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.

The Measurement of Boundary Layers on a Compressor Blade in Cascade: Part 2—Suction Surface Boundary Layers

1988 ◽  
Vol 110 (1) ◽  
pp. 138-145 ◽  
Author(s):  
S. Deutsch ◽  
W. C. Zierke
Keyword(s):  
Boundary Layers ◽  
Separation Bubble ◽  
Incidence Angle ◽  
Outer Region ◽  
Leading Edge ◽  
Compressor Blade ◽  
Surface Boundary ◽  
Suction Surface ◽  
Circular Arc ◽  
Edge Separation

Using the facility described in Part 1 [29], eleven detailed velocity and turbulence intensity profiles are obtained on the suction surface of a double circular arc blade in cascade. At the measured incidence angle of 5 deg, transition through a leading edge separation bubble occurs before 2.6 percent chord. A continuing recovery from this leading edge separation is apparent in the measured boundary layer profiles at 2.6 and 7.6 percent chord. Recovery appears to be complete by 12.7 percent chord. The data then illustrate the evolution of the nonequilibrium turbulent boundary layers as they approach a second region of separation. Following the criteria established by Simpson et al. [1], we find that intermittent separation occurs near 60 percent chord while detachment occurs at 84.2 percent chord. Comparison between the measured profiles and the sublimation visualization studies indicates that the flow visualization is signaling the location of incipient detachment (1 percent instantaneous backflow). Measured profiles are also considered in light of similarity techniques for boundary layers approaching separation. Outer region similarity is shown to vanish for profiles downstream of detachment.

Loss Production in the Wake of a Simulated Subsonic Turbine Blade

1996 ◽  
Author(s):  
Quentin D. Roberts ◽  
John D. Denton
Keyword(s):  
Boundary Layer ◽  
Turbine Blade ◽  
Vortex Street ◽  
Pressure Probe ◽  
Surface Boundary ◽  
Suction Surface ◽  
Surface Boundary Layer ◽  
Loss Coefficients ◽  
Streamwise Distance ◽  
Profile Loss

The time-averaged flow in the wake of a model of a turbine blade was surveyed using a three-hole pressure probe; boundary layer traverses were also carried out using a flattened pilot probe. Loss coefficients were derived from the mass-weighted deficit in stagnation pressure. Results showing the progression of loss with streamwise distance along the surface and the wake of the model are presented. It was found that the loss generated in the wake comprised one-third of the profile loss when a well-developed vortex street was present in the wake. This proportion was reduced by increasing the thickness of the suction surface boundary layer, and by simulating the deviation that occurs in a real turbine blade. In both cases the strength of the vortex street was also shown to have been reduced.

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.

The Measurement of Boundary Layers on a Compressor Blade in Cascade: Part 4 — Flow Fields for Incidence Angles of −1.5 and −8.5 Degrees

1989 ◽  
Author(s):  
W. C. Zierke ◽  
S. Deutsch
Keyword(s):  
Boundary Layers ◽  
Separation Bubble ◽  
Measurement Techniques ◽  
Leading Edge ◽  
Compressor Blade ◽  
Laser Doppler ◽  
Laser Doppler Velocimeter ◽  
Suction Surface ◽  
Laminar Separation ◽  
Pressure Surface

Measurements, made with laser Doppler velocimetry, about a double-circular-arc compressor blade in cascade are presented for −1.5 and −8.5 degree incidence angles and a chord Reynolds number near 500,000. Comparisons between the results of the current study and those of our earlier work at a 5.0 degree incidence are made. It is found that in spite of the relative sophistication of the measurement techniques, transition on the pressure surface at the −1.5 degree incidence is dominated by a separation “bubble” too small to be detected by the laser Doppler velocimeter. The development of the boundary layers at −1.5 and 5.0 degrees are found to be similar. In contrast to the flow at these two incidence angles, the leading edge separation “bubble” is on the pressure surface for the −8.5 degree incidence. Here, all of the measured boundary layers on the pressure surface are turbulent — but extremely thin — while on the suction surface, a laminar separation/turbulent reattachment “bubble” lies between roughly 35% and 60% chord. This “bubble” is quite thin, and some problems in interpreting backflow data.

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.

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