Stall Inception in a Compressor with Subsonic, Transonic, and Supersonic Inlet Conditions

2022 ◽  
pp. 1-11
Author(s):  
Semi Kim ◽  
Jaeho Choi ◽  
Ewan Gunn ◽  
Tobias Brandvik ◽  
Young Seok Kang
Keyword(s):  
Stall Inception ◽  
Supersonic Inlet ◽  
Inlet Conditions

scholarly journals Interpreting Aerodynamics of a Transonic Impeller from Static Pressure Measurements

2018 ◽  
Vol 2018 ◽  
pp. 1-9
Author(s):  
Fangyuan Lou ◽  
John Charles Fabian ◽  
Nicole Leanne Key
Keyword(s):  
Flow Field ◽  
High Speed ◽  
Static Pressure ◽  
Pressure Ratio ◽  
Leading Edge ◽  
Pressure Measurements ◽  
High Loss ◽  
Supersonic Inlet ◽  
Mach Numbers ◽  
Inlet Conditions

This paper investigates the aerodynamics of a transonic impeller using static pressure measurements. The impeller is a high-speed, high-pressure-ratio wheel used in small gas turbine engines. The experiment was conducted on the single stage centrifugal compressor facility in the compressor research laboratory at Purdue University. Data were acquired from choke to near-surge at four different corrected speeds (Nc) from 80% to 100% design speed, which covers both subsonic and supersonic inlet conditions. Details of the impeller flow field are discussed using data acquired from both steady and time-resolved static pressure measurements along the impeller shroud. The flow field is compared at different loading conditions, from subsonic to supersonic inlet conditions. The impeller performance was strongly dependent on the inducer, where the majority of relative diffusion occurs. The inducer diffuses flow more efficiently for inlet tip relative Mach numbers close to unity, and the performance diminishes at other Mach numbers. Shock waves emerging upstream of the impeller leading edge were observed from 90% to 100% corrected speed, and they move towards the impeller trailing edge as the inlet tip relative Mach number increases. There is no shock wave present in the inducer at 80% corrected speed. However, a high-loss region near the inducer throat was observed at 80% corrected speed resulting in a lower impeller efficiency at subsonic inlet conditions.

Investigation of Stall Process in a Centrifugal Compressor With a Volute Under Transonic Conditions

2019 ◽  
Author(s):  
Ce Yang ◽  
Wenli Wang ◽  
Hanzhi Zhang ◽  
Yanzhao Li ◽  
Ding Tong ◽  
...  
Keyword(s):  
Flow Field ◽  
Centrifugal Compressor ◽  
Static Pressure ◽  
Cell Formation ◽  
Internal Flow ◽  
Growth Zone ◽  
Asymmetric Structure ◽  
Stall Inception ◽  
Stall Cells ◽  
Inlet Conditions

Abstract In a centrifugal compressor with a volute, the internal flow field is circumferentially nonuniform owing to the asymmetric structure of the volute. Currently, the mechanisms by which the volute influences the stall inception circumferential position and the stall process in a transonic centrifugal compressor are not clear. In this study, the stall process in the centrifugal compressor with a volute is investigated under transonic inlet conditions. Obtained by experimental and simulation results, the static pressure distributions around the casing wall are compared with each other. Thereafter, an unsteady simulation is conducted on the stall process under transonic inlet conditions. By analyzing the stall cell evolution pattern at the impeller inlet, the stall process can be divided into three stages: stall onset, stall development, and stall maturation. The spike-type stall inceptions occur twice at the tip in the circumferential 135° position of the impeller inlet. This circumferential position is also the affected position of the high static pressure region induced by the volute tongue. Because of the circumferentially nonuniform flow field, there is a stall cell decay zone and a stall cell formation/growth zone at the impeller inlet. For the compressor under study, the approximate circumferential range of 135° to 270° is the decay zone, and the circumferential range of 270° to 360° is the formation/growth zone. The stall inception cannot occur in the decay zone. However, the stall cells can pass through the decay zone when the stall cell size is large enough. The first stall inception cannot propagate circumferentially, while the second one can. The propagation speed of stall cells in the circumferential direction is at approximately 70% of the rotational speed of impeller.

Numerical Performance Assessment of a Turbine Blade Row Operating Under Pulsed Detonation Combustion Inlet Conditions

2019 ◽  
Author(s):  
Cleopatra Cuciumita ◽  
Christian Oliver Paschereit
Keyword(s):  
Gas Turbines ◽  
Practical Implementation ◽  
Pressure Losses ◽  
Thermodynamic Cycles ◽  
Supersonic Inlet ◽  
Pulsed Detonation ◽  
Detonation Combustion ◽  
Inlet Conditions ◽  
Low Efficiency ◽  
New Research

Abstract Pulsed detonation combustion is not a new research topic. However, since the detonation process was first observed in 1881, the interest in it grew substantially in the last decades. Because the gas turbines have reached their technological maturity, the scientific community has started looking into novel thermodynamic cycles, such as detonation-based cycles. Numerous studies have been recently published in the field of pulsed detonation combustion, both numerical and experimental and major breakthroughs have been achieved in understanding and controlling the phenomena. However, the topic remains of mostly academic interest, one of the reasons is that practical implementation of it is reliant turbomachinery that would efficiently convert the pulsed, high peak, pressure into useful work. The few studies conducted on classical, existing turbines, show an efficiency of around 50% when coupled with a PDC. The low efficiency has been directly connected with the shock wave losses. For this reason, the design of turbines with supersonic inlet, and associated performance assessments have been researched. This work, however, has supersonic steady inlet Mach number or sinusoidal pulsating conditions around an average subsonic value. No public literature exists on the performances of turbines operating at pulsating inlet conditions similar to the outlet of a PDC. The current paper tackles exactly this issue. The geometry for a turbine stator row was designed based on supersonic inlet design criteria. This geometry was then subjected to CFD numerical simulations. First, the pressure losses associated with a constant supersonic inlet were numerically determined to be a little over 26%. The next step was to assess the pressure losses of the same turbine row geometry in a transient approach. This time, the inlet conditions were set to be variable in time. The values were taken from a 1D in-house code computing the parameters at the outlet of a PDC working on hydrogen and air under stoichiometric conditions. This inlet conditions give a much better insight with respect to the flow within a turbine row when coupled with a PDC. It was observed that the pressure losses, computed as a time average for a period corresponding to the PDC functioning frequency were of 12%. This value is much less than that for a constant supersonic inlet, mostly due to the turbine being exposed to the shock waves for less time.

Investigation of Pre-stall Mode and Pip Inception in High-Speed Compressors Through the Use of Correlation Integral

1999 ◽  
Vol 121 (4) ◽  
pp. 743-750 ◽  
Author(s):  
Michelle M. Bright ◽  
Helen K. Qammar ◽  
Leizhen Wang
Keyword(s):  
Flow Field ◽  
High Speed ◽  
Length Scale ◽  
Flow Strength ◽  
Stall Inception ◽  
Spectral Techniques ◽  
Correlation Integral ◽  
Signal Activity ◽  
Inlet Conditions ◽  
Value Changes

Five high-speed compressor configurations are used to identify pre-stall pressure signal activity under clean and distorted inlet conditions, and under steady injection and controlled injection conditions. Through the use of a nonlinear statistic called correlation integral, variations in the compressor dynamics are identified from the pre-stall pressure activity far before variations (modal or pip) are observed visually in the wall static pressure measurements. The correlation integral not only discerns changing dynamics of these compressors prior to stall, but is now used to measure the strength of the tip flow field for these five high-speed compressors. Results show that correlation integral value changes dramatically when the stall inception is modal; and it changes less severely when the stall inception is through pip disturbances. This algorithm can therefore distinguish from the pre-stall pressure traces when a machine is more likely to stall due to pips versus modes. When used in this manner, the correlation integral thus provides a measure of tip flow strength. The algorithm requires no predisposition about the expected behavior of the data in order to detect changing dynamics in the compressor; thus, no pre-filtering is necessary. However, by band-pass filtering the data, one can monitor changing dynamics in the tip flow field for various frequency regimes. An outcome of this is to associate changes in correlation integral value directly with frequency specific events occurring in the compressor, i.e., blade length scale events versus long length scale acoustic events. The correlation integral provides a potential advantage over linear spectral techniques because a single sensor is used for detection and analysis of the instabilities.

scholarly journals Low pressure ratio transonic fan stall with radial distortion

2020 ◽  
Vol 4 ◽  
pp. 226-237
Author(s):  
Tim Williams ◽  
Cesare Hall ◽  
Mark Wilson
Keyword(s):  
High Speed ◽  
Pressure Ratio ◽  
Full Range ◽  
Perforated Plate ◽  
Low Pressure ◽  
Suction Surface ◽  
Stall Inception ◽  
Radial Distortion ◽  
Inlet Conditions ◽  
Transonic Fan

Numerical methods that can predict stall behaviour with non-uniform inlet conditions allow assessment of the stable operating range across flight conditions during the design of fan stages for civil aircraft. To extend the application of methods validated with clean inflow, the effect of a tip low radial distortion on the stall behaviour of a low pressure ratio transonic fan has been investigated using both high speed experiments and 3D URANS computations. The distortion is generated in the experiment using a perforated plate and this is fully represented within the computational mesh. This enables computations to reproduce the full range of flow conditions accurately without adjusting the inlet boundary condition. Both the calculations and measurements show that the presence of the distortion decreases the stall cell rotational speed and increases the cell circumferential extent. In the calculations, the cell speed reduced from 87% to 67% of shaft speed, compared to a change of 82% to 58% in the experiment. With and without distortion, the computations show how stall inception stems from blockage formed by flow separation from the tip-section suction surface, behind the shock. In the distorted case, the more forward shock position produces the blockage further upstream, causing a greater reduction of flow to adjacent passages. This leads to a stall cell in the distorted case that is around 80% larger.

scholarly journals Effect of inlet turbulence on compressible flow through the tip-section turbine blade cascade with supersonic inlet

2018 ◽  
Vol 168 ◽  
pp. 02007
Author(s):  
Petr Straka ◽  
Jaromír Příhoda ◽  
David Fenderl ◽  
Bartoloměj Rudas
Keyword(s):  
Compressible Flow ◽  
Turbine Blade ◽  
Reynolds Stress Model ◽  
Numerical Code ◽  
Bypass Transition ◽  
Free Stream Turbulence ◽  
Blade Cascade ◽  
Supersonic Inlet ◽  
Commercial Code ◽  
Inlet Conditions

The contribution deals with the numerical simulation of 2D compressible flow though the tip-section turbine blade cascade with the supersonic inlet boundary conditions. The simulation was carried out by the in-house numerical code using the explicit algebraic Reynolds stress model completed by the bypass transition model with the algebraic equation for the intermittency coefficient. The γ-Re model implemented in the commercial code Fluent was used for the comparison. Predictions carried out for the nominal conditions were focused on the effect of inlet free-stream turbulence on the flow structure in the blade cascade under supersonic inlet conditions. Numerical results were compared with experimental data.

Improved Blade Profile Loss and Deviation Angle Models for Advanced Transonic Compressor Bladings: Part II—A Model for Supersonic Flow

1996 ◽  
Vol 118 (1) ◽  
pp. 81-87 ◽  
Author(s):  
W. M. Ko¨nig ◽  
D. K. Hennecke ◽  
L. Fottner
Keyword(s):  
Present Report ◽  
Axial Flow ◽  
Deviation Angle ◽  
New Model ◽  
Transonic Compressor ◽  
Supersonic Inlet ◽  
Blade Row ◽  
Flow Situation ◽  
Inlet Conditions ◽  
Profile Loss

New blading concepts as used in modern transonic axial-flow compressors require improved loss and deviation angle correlations. The new model presented in this paper incorporates several elements and treats blade-row flows having subsonic and supersonic inlet conditions separately. The second part of the present report focuses on the extension of a well-known correlation for cascade losses at supersonic inlet flows. It was originally established for DCA bladings and is now modified to reflect the flow situation in blade rows having low-cambered, arbitrarily designed blades including precompression blades. Finally, the steady loss increase from subsonic to supersonic inlet-flow velocities demonstrates the matched performance of the different correlations of the new model.

Improved Blade Profile Loss and Deviation Angle Models for Advanced Transonic Compressor Bladings: Part II — A Model for Supersonic Flow

1994 ◽  
Author(s):  
W. M. König ◽  
D. K. Hennecke ◽  
L. Fottner
Keyword(s):  
Present Report ◽  
Axial Flow ◽  
Deviation Angle ◽  
New Model ◽  
Transonic Compressor ◽  
Supersonic Inlet ◽  
Blade Row ◽  
Flow Situation ◽  
Inlet Conditions ◽  
Profile Loss

New blading concepts as used in modern transonic axial-flow compressors require improved loss and deviation angle correlations. The new model presented in this paper incorporates several elements and treats separately blade-row flows having subsonic and supersonic inlet conditions. The second part of the present report focuses on the extension of a well-known correlation for cascade losses at supersonic inlet-flows. It was originally established for DCA-bladings and is now modified to reflect the flow situation in blade-rows having low-cambered, arbitrarily designed blades including precompression blades. Finally, the steady loss increase from subsonic to supersonic inlet-flow velocities demonstrates the matched performance of the different correlations of the new model.

A Rational Analysis of Periodic Flow Perturbation in Supersonic Two-Dimensional Cascade

1979 ◽  
Vol 101 (3) ◽  
pp. 431-439 ◽  
Author(s):  
Ron-Ho Ni
Keyword(s):  
Flow Equation ◽  
Two Dimensional ◽  
Rational Analysis ◽  
Axial Velocity Component ◽  
Flow Perturbation ◽  
Supersonic Inlet ◽  
Previous Solution ◽  
Inlet Conditions ◽  
Unsteady Loading ◽  
Physical Phenomena

An analytical formulation which can be applied to obtain unsteady aerodynamic solutions for a cascade of flat plate blades oscillating in subsonic or supersonic flow with either subsonic or supersonic axial velocity component is presented. In the analysis, the flow is assumed to be two-dimensional and isentropic and the blades are undergoing small amplitude harmonic oscillations. The method of superposition of basic wave solutions of the linearized unsteady flow equation is used to construct the flow field induced by the harmonic motion of blades in cascade. This method leads to an integral equation from which the unsteady loading on a blade can be determined. Since the equation applied to both subsonic and supersonic inlet conditions, the present approach provides a unified basis for analyzing and understanding the complex physical phenomena associated with flow past vibrating cascades. The technique used to determine a solution for an unsteady supersonic cascade is also described and the results obtained are shown to agree with those from previous solution.

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