1. School of Aerospace Engineering MITE Numerical Simulation of Centrifugal Compressor Stall and Surge Saeid NiaziAlex SteinLakshmi N. Sankar School of Aerospace Engineering Georgia Institute of Technology Supported by the U.S. Army Research Office Under the Multidisciplinary University Research Initiative (MURI) on Intelligent Turbine Engines

2. School of Aerospace Engineering MITE Overview Motivation and Objectives Mathematical Formulation –Boundary Conditions Simulation Setup Results –Design and Off-Design Condition Validations –Stall Control (Bleed Valve) Conclusions Future Work

3. School of Aerospace Engineering MITE Motivation and Objectives pp m. With Control Without Control Centrifugal Compressors Are Widely Used in Many Practical Applications (Rotorcraft, Turbomachinery, Tanks) Develop a Numerical Scheme to Model and Understand Compressor Stall & Surge. Explore Active & Passive Control Strategies to Extend Useful Operating Range of Compressor.

4. School of Aerospace Engineering MITE MATHEMATICAL FORMULATION    t q  dV  E ˆ i  F ˆ j  G ˆ k     ndS  R ˆ i  S ˆ j  T ˆ k     ndS Reynolds Averaged Navier-Stokes Equations in Finite Volume Representation: where, q is the state vector. E, F, and G are the inviscid fluxes, and R, S, and T are the viscous fluxes. A cell-vertex finite volume formulation using Roe’s scheme is used for the present simulations.

5. School of Aerospace Engineering MITE * * i-1 i i+1 i+2 Cell face i+1/2 Stencil for q left Stencil for q right Left Right MATHEMATICAL FORMULATION A four point stencil is used to compute the inviscid flux terms at the cell faces as shown below: This makes the scheme third-order accurate in space.

6. School of Aerospace Engineering MITE MATHEMATICAL FORMULATION The viscous fluxes are computed to second order spatial accuracy. A three-factor ADI scheme with second-order artificial damping on the LHS is used to advance the solution in time. The scheme is first-order accurate in time. The Spalart-Allmaras turbulence model is used in the present simulations.

7. School of Aerospace Engineering MITE NASA Low Speed Centrifugal Compressor 20 Full Blades with 55° Backsweep Inlet Diameter 0.87 m Exit Diameter 1.52 m Design Conditions: –Mass Flow Rate 30 kg/sec –1862 RPM –Total Pressure Ratio 1.14 SIMULATION SETUP

8. School of Aerospace Engineering MITE Single Passage Grid Modeling SIMULATION SETUP Grid Size: 65x31x21 = 42315 points

9. School of Aerospace Engineering MITE Inlet: p 0,T 0,v,w specified; Riemann- Invariant extrapolated from Interior Exit: p back specified; all other quantities extrapolated from Interior Solid Walls: no-slip velocity conditions; p &  extrapolated from Interior Periodic Boundaries: Properties are averaged on either side of the boundary SIMULATION SETUP Boundary Conditions

10. School of Aerospace Engineering MITE Blade Pressure at Different Span Stations RESULTS (Design Conditions ) Good Agreement Between CFD and Experiment

11. School of Aerospace Engineering MITE Blade Pressure at Different Span Stations RESULTS (Design Conditions ) Good Agreement Between CFD and Experiment

12. School of Aerospace Engineering MITE Blade Pressure at Different Span Stations RESULTS (Design Conditions ) Slight Difference Between CFD and Experiment Due to Tip Vortex

13. School of Aerospace Engineering MITE Shroud Pressure Distribution Near Blades RESULTS (Design Conditions )

14. School of Aerospace Engineering MITE Shroud Pressure Distribution RESULTS (Design Conditions ) Pressure Increases Along Blade Passage

15. School of Aerospace Engineering MITE RESULTS (Design & Off-Design ) Impeller Performance Characteristic

16. School of Aerospace Engineering MITE Relative Velocity (Colored by Pressure) RESULTS (Design Conditions ) Flow is Well Attached. Very Small Regions of Separation Occur Near Shroud Wall (Enlarged View)

17. School of Aerospace Engineering MITE Relative Velocity (Colored by Pressure) RESULTS (Design Conditions ) Diffuser Region is Well Behaved No Separation

18. School of Aerospace Engineering MITE Relative Velocity (Colored by Pressure) RESULTS (Off-Design Conditions ) Diffuser Region Shows Small Separation Onset of Instabilities

19. School of Aerospace Engineering MITE Relative Velocity (Colored by Pressure) RESULTS (Off-Design Conditions ) Diffuser Region Completely Separated Diffuser Stall Occurs

20. School of Aerospace Engineering MITE Comparison Between Design & Off-Design RESULTS (Design & Off-Design )

21. School of Aerospace Engineering MITE Relative Velocity Colored by Pressure RESULTS (Bleed Valve ) Bleeding Greatly Improves Flow Behavior Diffuser Stall is Suppressed

22. School of Aerospace Engineering MITE Effects of Diffuser Bleeding RESULTS (Bleed Valve )

23. School of Aerospace Engineering MITE Effects of Diffuser Bleeding RESULTS (Bleed Valve ) Bleeding Reduces Reverse Flow in Compressor Exit Plane

24. School of Aerospace Engineering MITE CONCLUSIONS A 3-D unsteady compressible flow solver for modeling centrifugal compressors has been developed and validated. Good agreement with experiments have been obtained for a Low Speed Centrifugal Compressor (LSCC) tested at NASA Lewis Research Center. For the LSCC instabilities were found to originate in the diffuser region. Stall and surge may be eliminated by the use of bleed valves on the diffuser walls.

25. School of Aerospace Engineering MITE Continue to work on control issues, e.g. unsteady bleeding, recirculation. Development and implementation of unsteady boundary conditions. Apply compressor code to centrifugal compressor with higher pressure ratios. FUTURE WORK