Case Study No: FTLCFD08
Introduction
The NASA High-Efficiency Centrifugal Compressor (HECC) stage was developed at NASA Glenn Research Center as a benchmark configuration for small gas turbine engines, especially for rotorcraft applications. Its compact design, high-pressure ratio per stage, and robustness against varying operating conditions make it a reference case for both experimental and numerical investigations. The HECC stage has been widely studied for its aerodynamic performance, focusing particularly on the effects of diffuser design and blade tip clearance on overall compressor efficiency, surge margin, and pressure recovery. CFD provides a valuable approach to understanding the detailed flow physics within the HECC stage, where leakage flows, secondary vortices, and diffuser interactions strongly affect performance.
The purpose of the study is to perform a CFD analysis on the HECC stage and validate the CFD results with experimental data by comparing the stage performance parameters.
Figure 1: HECC full model
Computational Domain
The HECC compressor stage comprises an inlet passage, an impeller with 15 main blades and splitter blades, a diffuser with 20 main and splitter vanes, and exit guide vanes made of 60 cascade blades positioned perpendicular to the diffuser exit. The impeller–shroud tip clearance is 0.30 mm (about 1.93% of the impeller exit blade height). To reduce computational cost, a sector model with periodic boundaries was used.
Figure 2: Computational domain of HECC sector model
Mesh
A structured multi-block mesh was generated using TurboGrid. A structured mesh with boundary layers was generated with H-Type grid in the main flow region and O-Type grid near the edges of the blades/vanes to improve the mesh quality.
Figure 3: Mesh details
Numerical Model
A 3D, steady state CFD analysis was performed using ANSYS CFX 2024R2. Air was used as the working fluid. The nominal speed of the compressor was 21,789 rpm and CFD analysis was done with available experimental data for different speedlines. The k-ω SST model was used for turbulence modelling. The impeller stage in the compressor was modeled using the multiple reference frame method to account for the impeller blade rotation. At the inlet, total temperature and total pressure boundary conditions were applied with standard atmospheric conditions. A static pressure outlet boundary condition was used to simulate conditions near the choke side of the speed line, while a mass flow outlet or corrected mass flow outlet boundary condition was applied for points near the surge side.
| Location | Boundary condition |
|---|---|
| Inlet | Total pressure and Total temperature |
| Outlet | Static pressure/Mass flow rate |
| Walls | No slip adiabatic walls |
| Impeller shroud | counter rotating wall |
| Rotor speed | 100%, 90% and 75% of nominal speed |
Table 1: Boundary conditions details
Results
The isentropic stage efficiency and pressure ratio were evaluated across a wide range of mass flow rates. While some discrepancies in pressure ratio and efficiency appeared near choke, the CFD results showed good agreement with the experiment as the operating point moved toward surge. Across all simulated speedlines, the variation in pressure ratio was within 1% of the experimental data, except at choke. Similarly, the stage efficiency deviation was less than 1% at the 75% and 90% speedlines, and within 3% at the 100% speedline, as shown in Fig. 4. Overall, the simulations predicted a choke mass flow rate approximately 4% higher than the measured value. Figure 5 and 6 represents the temperature variation and velocity field along with vectors respectively for the HECC stage.
Figure 4: Stage performance parameters, comparison CFD Vs Experimental
Conclusion
- Benchmarking for the HECC stage was done and the stage performance parameters were computed.
- The CFD data was validated with the available experimental data for different speeds and good agreement was observed between CFD and experimental data.
References
- Medic, Gorazd, et al. High efficiency centrifugal compressor for rotorcraft applications. No. NASA/CR-2014-218114. National Aeronautics and Space Administration, 2014.
- Harrison, Herbert M., Ezra O. McNichols, and Matthew R. Blaha. “NASA Small Engine Components Compressor Test Facility: High Efficiency Centrifugal Compressor Vaneless Diffuser and Transition Duct Configurations.” Turbo Expo: Power for Land, Sea, and Air. Vol. 87110. American Society of Mechanical Engineers, 2023.
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