The Unsteady Loss in One-Stage Transonic Compressor Under Peak Efficiency and Near Stall Conditions

2008 ◽  
Author(s):  
Yutao Sun ◽  
Yu-Xin Ren ◽  
Song Fu ◽  
Aspi R. Wadia
Keyword(s):  
Operating Conditions ◽  
Flow Structures ◽  
Surface Boundary ◽  
Peak Efficiency ◽  
Suction Surface ◽  
Loss Mechanism ◽  
Transonic Compressor ◽  
Thermodynamic Irreversibility ◽  
Unsteady Rans ◽  
Blade Row Interaction

The loss mechanism due to the blade row interaction effects in a compressor is an important and interesting problem. In the present paper, the loss in a compressor stage under both peak efficiency (PE) and near stall (NS) operating conditions is studied numerically by solving the 3D unsteady RANS equations. The loss is measured in terms of the thermodynamic irreversibility. In order to examine the relationship between loss and flow structures, an intensive parameter called the intensity of irreversibility is derived. The relative importance of various flow structures in loss production has been identified by the orders of the intensity of irreversibility. An audit of the irreversibility has been carried out, which shows the importance of suction surface boundary layers and the tip leakages in loss generation. Under the NS conditions, the tip leakages are the most important sources of loss production. The unsteady evolution of irreversibility in the rotor and stator passage is also analyzed.

scholarly journals Blade Row Interaction Effects on the Performance of a Moderately Loaded NASA Transonic Compressor Stage

2002 ◽  
Author(s):  
Dale E. Van Zante ◽  
Wai-Ming To ◽  
Jen-Ping Chen
Keyword(s):  
Experimental Data ◽  
Velocity Field ◽  
Interaction Effects ◽  
Operating Conditions ◽  
Laser Doppler Velocimeter ◽  
Compressor Stage ◽  
Peak Efficiency ◽  
Transonic Compressor ◽  
Blade Row ◽  
Blade Row Interaction

Blade row interaction effects on loss generation in compressors have received increased attention as compressor work-per-stage and blade loading have increased. Two dimensional Laser Doppler Velocimeter measurements of the velocity field in a NASA transonic compressor stage show the magnitude of interactions in the velocity field at the peak efficiency and near stall operating conditions. The experimental data are presented along with an assessment of the velocity field interactions. In the present study the experimental data are used to confirm the fidelity of a three-dimensional, time-accurate, Navier Stokes calculation of the stage using the MSU-TURBO code at the peak efficiency and near stall operating conditions. The simulations are used to quantify the loss generation associated with interaction phenomena. At the design point the stator pressure field has minimal effect of the rotor performance. The rotor wakes do have an impact on loss production in the stator passage at both operating conditions. A method for determining the potential importance of blade row interactions on performance is presented.

Control of Shock Structure and Secondary Flow Field Inside Transonic Compressor Rotors Through Aerodynamic Sweep

1998 ◽  
Author(s):  
C. Hah ◽  
S. L. Puterbaugh ◽  
A. R. Wadia
Keyword(s):  
Aspect Ratio ◽  
Shock Structure ◽  
Secondary Flows ◽  
Operating Conditions ◽  
Numerical Analyses ◽  
Peak Efficiency ◽  
Stall Margin ◽  
Transonic Compressor ◽  
Low Aspect Ratio ◽  
Through Flow

The present paper reports a numerical study on the effects of aerodynamic sweep applied to a low-aspect-ratio, high-through-flow, state-of-the-art, axial transonic compressor design. Numerical analyses based on the Reynolds-averaged Navier-Stokes equations were used to obtain the performance of a conventional unswept rotor, a forward swept rotor, and an aft-swept rotor, at both design and off-design operating conditions. The numerical analyses predicted that the forward-swept rotor has a higher peak efficiency and a substantially larger stall margin than the baseline unswept rotor, and that the aft-swept rotor has a similar peak efficiency as the unswept rotor with a significantly smaller stall margin. The rig test confirmed the numerical assessment of the effects of aerodynamic sweep on the low-aspect-ratio, high-through-flow, transonic compressor rotor. Detailed analyses of the measured and calculated flow fields indicate that two mechanisms are primarily responsible for the differences in aerodynamic performance among these rotors. The first mechanism is a change in the radial shape of the passage shock near the casing by the endwall effect, and the second is the radial migration of low-momentum fluid to the blade tip region. Aerodynamic sweep can be used to control the shock structure near the endwall and the migration of secondary flows and, consequently, flow structures near the tip area for improved performance.

The Effects of Blade Loading on Trailing Edge Vortex Formation on a Highly Loaded Stator Upstream of a Transonic Rotor

2011 ◽  
Author(s):  
Kenneth P. Clark ◽  
Steven E. Gorrell
Keyword(s):  
Vortex Shedding ◽  
Operating Conditions ◽  
Strength Increase ◽  
Shock Strength ◽  
Suction Surface ◽  
Low Velocity ◽  
Trailing Edge ◽  
Blade Loading ◽  
Transonic Compressor ◽  
Velocity Region

Multiple high-fidelity time-accurate computational fluid dynamics simulations were performed to investigate the effects of upstream stator loading and rotor shock strength on vortex shedding characteristics in a single stage transonic compressor. Three loadings on the upstream stator row of decreased, nominal, and increased were studied. The time-accurate URANS code TURBO was used to generate periodic, quarter annulus simulations of the Blade Row Interaction compressor rig. It was observed that vortex shedding was synchronized to the passing of a rotor bow shock. Results show that vortex size and strength increase with stator loading. “Normal” and “large” shock-induced vortices formed on the stator trailing edge immediately after the shock passing, but the “large” vortices were strengthened at the trailing edge due to a low velocity region at the suction surface. This low velocity region was generated upstream on the suction surface from a shock-induced thickening of the boundary layer or separation bubble. The circulation of the “large” vortices was greater than the “normal” vortices by a factor of 1.7, 1.9 and 2.1 for decreased, nominal and increased deswirler loadings. At decreased loading only 24% of the measured vortices were considered “large” while at nominal loading 58% were “large”. An understanding of the unsteady interactions associated with blade loading and rotor shock strength in transonic stages will help compressor designers account for unsteady flow physics at design and off-design operating conditions.

Analysis and Prediction of Shock-Induced Vortex Circulation in Transonic Compressors

2015 ◽  
Vol 137 (12) ◽  
Author(s):  
Kenneth P. Clark ◽  
Steven E. Gorrell
Keyword(s):  
Vortex Shedding ◽  
Separation Bubble ◽  
Operating Conditions ◽  
Bow Shock ◽  
Shock Strength ◽  
Suction Surface ◽  
Low Velocity ◽  
Transonic Compressor ◽  
Dynamics Simulations ◽  
Velocity Region

Multiple high-fidelity time-accurate computational fluid dynamics simulations were performed to investigate the effects of upstream stator loading and rotor shock strength on vortex shedding characteristics in a single-stage transonic compressor. Three loadings on the upstream stator row of decreased, nominal, and increased loading in conjunction with three axial spacings of close, mid, and far were studied for this analysis. The time-accurate urans code turbo was used to generate periodic, quarter annulus simulations of the blade row interaction (BRI) compressor rig. It was observed that vortex shedding was synchronized to the passing of a rotor bow shock. Results show that vortex strength increases linearly with stator loading and rotor bow shock strength. “Normal” and “large” shock-induced vortices formed on the stator trailing edge (TE) immediately after the shock passing, but the large vortices were strengthened at the TE due to a low-velocity region on the suction surface. This low-velocity region was generated upstream on the suction surface from a shock-induced thickening of the boundary layer or separation bubble. The circulation of the large vortices was greater than the normal vortices by a factor of 1.7, 1.83, and 2.04 for decreased, nominal, and increased deswirler loadings. At decreased loading, only 24% of the measured vortices were considered large, while at nominal loading 58% were large. A model was developed to predict shock-induced vortex circulation from a known rotor bow shock strength, stator diffusion factor, and near-wake parameters. The model predicts the average vortex circulation very well, with 5% difference between predicted and measured values. An understanding of the unsteady interactions associated with blade loading and rotor shock strength in transonic stages will help compressor designers account for unsteady flow physics at design and off-design operating conditions.

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.

Reducing Low-Pressure Turbine Stage Blade Count Using Vortex Generator Jet Separation Control

2002 ◽  
Author(s):  
Rolf Sondergaard ◽  
Jeffrey P. Bons ◽  
Matthew Sucher ◽  
Richard B. Rivir
Keyword(s):  
Boundary Layer ◽  
Vortex Generator ◽  
Separation Control ◽  
Surface Boundary ◽  
Freestream Turbulence ◽  
Skew Angle ◽  
Suction Surface ◽  
Simple Boundary ◽  
Design Value ◽  
Vortex Generator Jets

An experimental investigation has been conducted into the feasibility of increasing blade spacing (pitch) at constant chord in a linear turbine cascade. Vortex generator jets (VGJs) located on the suction surface of each blade in the cascade are employed to maintain attached boundary layers despite the increasing tendency to separate due to the increased uncovered turning. Tests were performed at low Mach numbers and at blade Reynolds numbers between 25,000 and 75,000 (based on axial chord and inlet velocity). The vortex generator jets (30 degree injection angle and 90 degree skew angle) were operated with steady flow with momentum blowing ratios between zero and five, and from two spanwise rows of holes located at 45% and 63% axial chord. In the absence of control, pitch-averaged wake losses increase up to 600% as the blade pitch is increased from its design value to twice the design value. With the application of VGJs, these losses were driven down to or below the losses at the design pitch. The effectiveness of VGJs was found to increase modestly with increasing Reynolds number up to the highest value tested, Re = 75,000. The fluid phenomenon responsible for this remarkable range of effectiveness is clearly more than a simple boundary layer transition effect, as boundary layer trips installed on the same blades without VGJ blowing had no beneficial effect on blade losses. Also, tests conducted at elevated levels of freestream turbulence (4% at the cascade inlet) where the suction surface boundary layer is generally turbulent, showed wake loss reduction comparable to tests conducted at the nominal 1% freestream turbulence. For all configurations, blowing from the upstream row had the greatest wake influence. These findings open the possibility that future LPT designs could take advantage of active separation control using integrated VGJs to reduce the turbine part count and stage weight without significant increase in pressure losses.

scholarly journals The Design and Evaluation of a 14 Stage, 16:1 Pressure Ratio Compressor for an Industrial Gas Turbine

1996 ◽  
Author(s):  
Mohammad R. Saadatmand
Keyword(s):  
Experimental Data ◽  
Gas Turbine ◽  
Rotor Blade ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Operating Conditions ◽  
Suction Surface ◽  
Design Intent ◽  
Flow Through ◽  
Industrial Gas Turbine

The aerodynamic design process leading to the production configuration of a 14 stage, 16:1 pressure ratio compressor for the Taurus 70 gas turbine is described. The performance of the compressor is measured and compared to the design intent. Overall compressor performance at the design condition was found to be close to design intent. Flow profiles measured by vane mounted instrumentation are presented and discussed. The flow through the first rotor blade has been modeled at different operating conditions using the Dawes (1987) three-dimensional viscous code and the results are compared to the experimental data. The CFD prediction agreed well with the experimental data across the blade span, including the pile up of the boundary layer on the corner of the hub and the suction surface. The rotor blade was also analyzed with different grid refinement and the results were compared with the test data.

Impact of Multi-Row Aerodynamic Interaction on the Forced Response Behaviour of an Embedded Compressor Rotor

2020 ◽  
Author(s):  
Shreyas Hegde ◽  
Robert Kielb ◽  
Laith Zori ◽  
Rubens Campregher
Keyword(s):  
Higher Order ◽  
Forced Response ◽  
Operating Conditions ◽  
High Loading ◽  
Response Behavior ◽  
Wave Reflections ◽  
Peak Efficiency ◽  
Compressor Rotor ◽  
Higher Order Modes ◽  
Modal Force

Abstract This paper focuses on the impact of multi-row interaction on the forced response behavior of an embedded compressor rotor at higher order modes. The authors in previous papers have discussed about the multi-row influence at the torsional mode resonant crossing and this paper extends the study to higher order modes. The paper talks about both the steady and unsteady influence of having additional rows in the configuration. It makes use of the time transformation (TT) method available in CFX to reduce the number of passages required in each row. Since the number of vanes from both the stators and the inlet guide vanes (IGV) is the same, the excitations from upstream rows and the potential field influence of the downstream row all contribute to the forcing, which is quantified both in terms of modal force and individual blade response. This paper describes the multi-row influence on the chordwise bending modes at both the peak efficiency (PE) and the high loading (HL) operating condition. To ascertain this influence, a 3-row case with just the two neighboring stators (S1, R2, S2 a 4-row case with the downstream rotor as well (S1, R2, S2, R3) and a 5-row with the upstream IGV were considered. While the 3-row case helped to determine the influence of neighboring stators on the forcing, the 4-row case provided the influence of the downstream rotor on the forced response behavior. Since the number of IGV vanes was the same as the neighboring stators the nature of interference between the stator and IGV wakes was determined as well. The 4-row case helped investigate physical wave reflections off a downstream rotating row, which had a significant influence on the modal force. The final section of the paper focuses on the mistuning response, which essentially couples frequency variations with the structural and aerodynamic aspects to predict individual blade responses, which are compared to experimental data. A mistuning analysis was carried out with the frequency mistuning present in the experimental facility Some of the key conclusions from this investigation are: 1) The interference of the IGV with the downstream stator (S1) is destructive at peak efficiency and constructive at high loading in line with the previous observation at torsional modes; 2) Physical wave reflections are constructive at all operating conditions at higher order modes unlike torsional modes where it was destructive; 3) The 3-row case gives the most accurate prediction in terms of average blade response and the 5-row case in terms of maximum blade response. Hence one of the significant findings is that, the aeromechanical behavior can be ascertained to a great deal of accuracy using just 3-rows at higher order modes crossings.

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.

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