Numerical Studies of Stall and Surge Alleviation in Compressors Alex Stein, Saeid Niazi, and Lakshmi N. Sankar School of Aerospace Engineering Georgia Institute of Technology http://www.ae.gatech.edu/~lsankar/MURI Supported by the U.S. Army Research Office Under the Multidisciplinary University Research Initiative (MURI) on Intelligent Turbine Engines

Overview Objectives and Motivation Rotating Stall and Surge Flow Solver and Boundary Conditions DLR High-Speed Centrifugal Compressor Unsteady Surge Simulations Surge Control Using Air-Injection NASA Axial Rotor 67 Results Peak Efficiency Conditions Off-design Conditions Bleed Valve Control Conclusions

Objectives and Motivation Lines of Constant Rotational Speed Efficiency Choke Limit Surge Limit Flow Rate Total Pressure Rise Desired Extension of Operating Range Develop a numerical scheme to model and understand compressor stall and surge. Explore active and passive control strategies (Bleed Valve, Air-Injection) to extend useful operating range of compressors.

Motivation and Objectives Compressor instabilities can cause fatigue and damage to entire engine

Rotating Stall Rotating stall is a 2-D unsteady local phenomenon Types of rotating stall: Part-span Full-span 2 2 2 1 1 1 Blade 1 sees a high a. Blade 1 recovers. Blase 2 stalls. Blade 1 stalls.

Surge Mild Surge Deep Surge Mean Operating Point Pressure Rise Flow Rate Mean Operating Point Pressure Rise Flow Rate Peak Performance Limit Cycle Oscillations Time Flow Rate Period of Mild Surge Cycle Time Flow Rate Period of Deep Surge Cycle Flow Reversal

How to Control Instabilities Bleed Valves Diffuser bleed valves Pinsley, Greitzer, Epstein (MIT) Prasad, Neumeier, Haddad (GT) Movable plenum walls Gysling, Greitzer, Epstein (MIT) Guide vanes Dussourd (Ingersoll-Rand Research Inc.) Air-injection Murray (CalTech) Fleeter, Lawless (Purdue) Weigl, Paduano, Bright (MIT & NASA Lewis) Movable Plenum Walls Guide Vanes Air-Injection

GTTURBO3D Flow Solver Reynolds averaged Navier-Stokes equations in finite volume representation. A Four Point Stencil is used to compute the inviscid flux terms at the cell faces according to Roe’s formulation (Third-order accurate in space, first- or second-order accurate in time) The viscous fluxes are computed to second order spatial accuracy. Turbulence is modeled by one-equation Spalart-Allmaras model Code can handle multiple computational blocks and inlet distortions

Boundary Conditions (GTTURBO3D) Outflow boundary (coupling with plenum) Periodic Boundary at compressor inlet Solid Wall Boundary at compressor casing Periodic Boundary at diffuser at impeller blades at clearance gap at compressor hub Inflow Zonal Boundary

Outflow BC (GTTURBO3D) Conservation of mass: CFD Outflow Boundary Plenum Chamber u(x,y,z) = 0 pp(x,y,z) = const. isentropic ap, Vp mc . mt CFD Outflow Boundary Conservation of mass:

DLR High-Speed Centrifugal Compressor 40cm Designed and tested by DLR (Germany) High pressure ratio AGARD test case

DLR High-Speed Centrifugal Compressor 24 main blades 30 backsweep CFD-grid 141 x 49 x 33 (230,000 grid-points) Design Conditions: 22,360 RPM Mass flow = 4.0 kg/s Total pressure ratio = 4.7 Adiab. efficiency = 83% Exit tip speed = 468 m/s Inlet Mrel = 0.92

DLR-High-Speed-Results (Design Conditions) Static Pressure Along Shroud Excellent agreement between CFD and experiment

DLR-High-Speed-Results (Off-Design Conditions) Performance Characteristic Map Unsteady fluctuations are denoted by size of circles Design Fluctuations at 3.1 kg/sec are 30 times larger than at 4.6 kg/sec Surge Choke

DLR-High-Speed-Results (Surge Conditions) wt/2p, Mild surge cycles develop Surge amplitude grows to 60% of mean flow rate Surge frequency = 94 Hz (1/100 of blade passing frequency)

DLR-High-Speed-Results (Air-Injection-Setup) 0.04RInlet Casing 5° Rotation Axis Impeller RInlet Injection angle,  = 5º 3 to 6% injected mass flow rate

DLR-High-Speed-Results (Air-Injection) Different yaw angles, 3% injected mass flow rate Yaw angle directly affects the unsteady leading edge vortex shedding

DLR-High-Speed-Results (Air-Injection) Different yaw angles, 3% injected mass flow rate Optimum: Surge amplitude/main flow = 8 % Injected flow/main flow = 3.2 % Yaw angle = 7.5 degrees

DLR-High-Speed-Results (Air-Injection) Yaw angle vs DLR-High-Speed-Results (Air-Injection) Yaw angle vs. angle of attack, 3% injected mass flow rate Injection yaw angle directly affects leading edge angle of attack => maximum control for designer

DLR-High-Speed-Results (Air-Injection) Angle of attack is directly altered by injection Optimum injection yaw angle of 7.5 degrees yields best result

Axial Compressor (NASA Rotor 67) 22 Full Blades Inlet Tip Diameter 0.514 m Exit Tip Diameter 0.485 m Tip Clearance 0.61 mm Design Conditions: Mass Flow Rate 33.25 kg/sec Rotational Speed 16043 RPM (267.4 Hz) Rotor Tip Speed 429 m/sec Inlet Tip Relative Mach Number 1.38 Total Pressure Ratio 1.63 Adiabatic Efficiency 0.93 Hub 4 Blocks 73X32X21 Total of 196,224 cells

Literature Survey of NASA Rotor 67 Computation of the stable part of the design speed operating line: NASA Glenn Research Center (Chima, Wood, Adamczyk, Reid, and Hah) MIT (Greitzer, and Tan) U.S. Army Propulsion Laboratory (Pierzga) Alison Gas Turbine Division (Crook) University of Florence, Italy (Arnone ) Honda R&D Co., Japan (Arima) Effects of tip clearance gap: NASA Glenn Research Center (Chima and Adamczyk) MIT (Greitzer) Shock boundary layer interaction and wake development: NASA Glenn Research Center (Hah and Reid). End-wall and casing treatment: NASA Glenn Research Center (Adamczyk)

Relative Mach Contours at Mid-Span (Peak Efficiency)  IV III II I LE TE Spatially uniform flow at design conditions

Relative Mach Number at 90% Radius (Peak Efficiency) 30% Pitch 50% Pitch TE LE TE LE

Shock-Boundary Layer Interaction (Peak Efficiency) LE TE Shock Near Suction Side

Velocity Profile at Mid-Passage (Peak efficiency) TE Shock % Mass Flow rate Fluctuations % Pressure Fluctuations Fluctuations are very small (2%) Flow is well aligned. Very small regions of separation observed in the tip clearance gap(Enlarged view)

Enlarged View of Velocity Profile in the Clearance Gap (Peak efficiency) TE Clearance Gap The reversed flow in the gap and the leading edge vorticity grow in size and magnitude as the compressor operates at off-design conditions

Performance Map (NASA Rotor 67) measured mass flow rate at choke: 34.96 kg/s CFD choke mass flow rate: 34.76 kg/s C B D Unstable Conditions A Near Stall Controlled Peak Efficiency

Location of the Probes for Observing the Pressure and Velocity Fluctuations The probes are located at 30% chord upstream of the rotor and 90% span. They are fixed in space. LE TE  I II III IV I II III IV

Onset of the Stall (Clean Inlet) Wt/2p I II III IV I II III IV  Probes show identical fluctuations. Flow while unsteady, is still symmetric from blade to blade.

Onset of the Stall (Disturbed Inlet) II III IV  Wt/2p Inlet distortion simulated by dropping the stagnation pressure in one block by 20%. Flow is no longer symmetric from blade to blade. Frequency of rotating stall is NW/3.5, where NW : blade passage frequency.

Bleed Valve Control Un = mb/(rAb) Pressure, density and tangential One Tip Chord Pressure, density and tangential velocities are extrapolated from interior. . Un = mb/(rAb) Shroud Hub

% Mass Flow Rate Fluctuations Bleed Valve Control Without Bleed Valve With Bleed Valve % Total Pressure Fluctuations 3% bleed air reduces the total pressure fluctuations by 75% % Mass Flow Rate Fluctuations

Bleed Valve Control Axial Velocity Near LE After 1.5 Rev. % From Hub After 0.5 Rev. Bleed Valve.

Conclusions A 3-D numerical flow solver has been developed to investigate compressor instabilities. The flow solver has been applied to obtain a detailed understanding of surge and rotating stall phenomena in axial and centrifugal compressors. Air-injection and bleeding have been numerically analyzed as compressor control schemes. Surge margin extension was achieved for both compression systems. The proper application of air-injection is sensitive to the injection-parameters (e.g. yaw angle).