Randall M. Chriss

The Influence of Compressor Blade Row Interaction Modeling on Performance Estimates From Time-Accurate, Multistage, Navier–Stokes Simulations

The time-accurate, multi-stage, Navier-Stokes, turbomachinery solver TURBO was used to calculate the aero performance of a 2 1/2 stage, highly-loaded, high-speed, axial compressor. The goals of the research project were to demonstrate completion times for multi-stage, time-accurate simulations that are consistent with inclusion in the design process, and to assess the influence of differing approaches to modeling the effects of blade row interactions on aero performance estimates. Three different simulation setups were used to model blade row interactions: 1.) single passage per blade row with phase lag boundaries, 2.) multiple passages per blade row with phase lag boundaries, and 3.) a periodic sector (1/2 annulus sector). The simulations used identical inlet and exit boundary conditions and identical meshes. To add more blade passages to the domain, the single passage meshes were copied and rotated. This removed any issues of differing mesh topology or mesh density from the following results. The 1/2 annulus simulation utilizing periodic boundary conditions required an order of magnitude less iterations to converge when all three simulations were converged to the same level as assessed by monitoring changes in overall adiabatic efficiency. When using phase lag boundary conditions the need to converge the time history information necessitates more iterations to obtain the same convergence level. In addition to convergence differences, the three simulations gave different overall performance estimates where the 1/2 annulus case was 1.0 point lower in adiabatic efficiency than the single passage phase lag case. The interaction between blade rows in the same frame of reference set up spatial variations of properties in the circumferential direction which are stationary in that reference frame. The phase lag boundary condition formulation will not capture this effect because the blade rows are not moving relative to each other. Thus for simulations of more than two blade rows and strong interactions, a periodic simulation is necessary to estimate the correct aero performance. [Keywords: blade row interaction, numerical modeling, multi-stage compressor]

Upstream Wake Influences on the Measured Performance of a Transonic Compressor Stage

The ine uence of an upstream wake on the performance of a downstream compressor stage with transonic inlet conditionsisstudied.Experimentalresultsfrome xedplanecompressorexitinstrumentationshowthatdeepwakes, representative of heavily loaded stator wakes, persist further downstream than anticipated. The ine uences of the upstreamwakesarealiasedintoatypicalstatorpitche eld,thusmagnifyingtheir“ true” ine uence.Resultsalsoshow that reducing the axial spacing between the upstream stator blade-row and downstream rotor blade-row reduced the overall performance suggesting that mechanisms other than wake recovery are present for this compressor. This change in performance was determined to be associated with the upstream stator wakes. The work presented herein shows that the stator/rotor interaction is signie cant and should be accounted for in the design, testing, and analysis of transonic axial compressors.

Variations in Upstream Vane Loading With Changes in Back Pressure in a Transonic Compressor

Dynamic loading of an inlet guide vane (IGV) in a transonic compressor is characterized by unsteady IGV surface pressures. These pressure data were acquired for two spanwise locations at a 105 percent speed operating condition, which produces supersonic relative Mach numbers over the majority of the rotor blade span. The back pressure of the compressor was varied to determine the effects from such changes. Strong bow shock interaction was evident in both experimental and computational results. Variations in the back pressure have significant influence on the magnitude and phase of the upstream pressure fluctuations. The largest unsteady surface pressure magnitude, 40 kPa, was obtained for the near-stall mass flow condition at 75 percent span and 95 percent chord. Radial variation effects caused by the spanwise variation in relative Mach number were measured. Comparisons to a two-dimensional nonlinear unsteady blade/vane Navier-Stokes analysis show good agreement for the 50 percent span results in terms of IGV unsteady surface pressure. The results of the study indicate that significant nonlinear bow shock influences exist on the IGV trailing edge due to the downstream rotor shock system.

The Effects of Blade-Row Spacing on the Flow Capacity of a Transonic Rotor

Results from the ongoing Air Force Stage Matching Investigation are presented. In the present work the effect of upstream blade row wakes on the flow capacity of a downstream stage (where the rotor sets the choking flow) is investigated. An embedded stage was simulated by placing a set of wake generators (similar to inlet guide vanes) in front of a highly loaded single stage transonic core compressor. The wake generator-to-rotor axial spacing was varied in addition to the vane count. A complete parametric test matrix was completed in order to determine which parameters were important to the choking flow capacity.The results show that for axial spacings above 50% of the upstream axial blade chord, simple wake mixing alone fully accounts for the upstream flow losses and a simple mass flow rate correction based on the rotor face mass averaged total pressure is sufficient. At spacings closer than this, other effects or loss mechanisms may be present. If these loss sources do exist, they are of unknown origin and magnitude and so an embedded overflow condition for these spacings cannot be ruled out. Research is ongoing that will attempt to identify and clarify the relevant details associated with these close blade row spacings.Copyright

UNSTEADY BLADE ROW POTENTIAL INTERACTION IN A COMPRESSION STAGE

A set of inlet guide vane (IGV) unsteady surface pressure measurements is presented. The unsteady aerodynamic effects of a highly loaded, high speed downstream compression stage on the upstream inlet guide vane/stator surface pressures at a part speed operation is characterized by experimental and computational analysis methods. The axial spacing between the IGV and rotor was varied between 12%, 26%, and 56% of the rotor chord for a 70% speed near choke operating condition, which is subsonic. Unsteady IGV surface pressures were acquired for two spanwise locations on both blade surfaces. Significant potential interaction is evident up to the 50% chord location. The upstream potential effect is strongly nonlinear and three dimensional in character, even though the IGV flow is largely two dimensional. The higher harmonic characteristics of the upstream potential effect decreases as the spacing between the IGV and rotor is increased. Comparisons to a nonlinear unsteady multi-blade row Navier-Stokes analysis show a good trendwise agreement in the IGV unsteady surface pressure envelope results, however computational capabilities to predict the frequency character requires further study.

Experimental and Computational Results From the Nasa Lewis Low-Speed Centrifugal Impeller at Design and Part Flow Conditions

The NASA Lewis Low-Speed Centrifugal Compressor (LSCC) has been investigated with laser anemometry and computational analysis at two flow conditions: the design condition as well as a lower mass flow condition. Previously reported experimental and computational results at the design condition are in the literature (Hathaway et al. 1993). In that paper extensive analysis showed that inducer blade boundary layers are centrifuged outward and entrained into the tip clearance flow and hence contribute significantly to the throughflow wake. In this report results are presented for a lower mass flow condition along with further results from the design case.The data set contained herein consists of three-dimensional laser velocimeter results upstream, inside and downstream of the impeller. In many locations data have been obtained in the blade and endwall boundary layers. The data are presented in the form of throughflow velocity contours as well as secondary flow vectors.The results reported herein illustrate the effects of flow rate on the development of the throughflow momentum wake as well as on the secondary flow. The computational results presented confirm the ability of modern computational tools to accurately model the complex flow in a subsonic centrifugal compressor. However, the blade tip shape and tip clearance must be known in order to properly simulate the flow physics. In addition, the ability to predict changes in the throughflow wake, which is largely fed by the tip clearance flow, as the impeller is throttled should give designers much better confidence in using computational tools to improve impeller performance.© 1994 ASME

The Influence of Compressor Blade Row Interaction Modeling on Performance Estimates From Time-Accurate, Multi-Stage, Navier-Stokes Simulations

The time-accurate, multi-stage, Navier-Stokes, turbomachinery solver TURBO was used to calculate the aero performance of a 2 1/2 stage, highly-loaded, high-speed, axial compressor. The goals of the research project were to demonstrate completion times for multi-stage, time-accurate simulations that are consistent with inclusion in the design process, and to assess the influence of differing approaches to modeling the effects of blade row interactions on aero performance estimates. Three different simulation setups were used to model blade row interactions: 1.) single passage per blade row with phase lag boundaries, 2.) multiple passages per blade row with phase lag boundaries, and 3.) a periodic sector (1/2 annulus sector). The simulations used identical inlet and exit boundary conditions and identical meshes. To add more blade passages to the domain, the single passage meshes were copied and rotated. This removed any issues of differing mesh topology or mesh density from the following results. The 1/2 annulus simulation utilizing periodic boundary conditions required an order of magnitude less iterations to converge when all three simulations were converged to the same level as assessed by monitoring changes in overall adiabatic efficiency. When using phase lag boundary conditions the need to converge the time history information necessitates more iterations to obtain the same convergence level. In addition to convergence differences, the three simulations gave different overall performance estimates where the 1/2 annulus case was 1.0 point lower in adiabatic efficiency than the single passage phase lag case. The interaction between blade rows in the same frame of reference set up spatial variations of properties in the circumferential direction which are stationary in that reference frame. The phase lag boundary condition formulation will not capture this effect because the blade rows are not moving relative to each other. Thus for simulations of more than two blade rows and strong interactions, a periodic simulation is necessary to estimate the correct aero performance.Copyright

Numerical Study of Embedded Supersonic Compressor Stages

A numerical case study of a multistage, highly-loaded, relative supersonic compressor is presented. The purpose of the investigation was to highlight the changing shock structure while throttling the compressor and to give insight into possible compressor instabilities. The computational fluid dynamic (CFD) study was conducted with the NASA code ADPAC, utilizing the mixing-plane assumption for the boundary condition between adjacent, relatively-rotating blade rows. A steady, five-blade-row, numerical simulation using the Baldwin-Lomax turbulence model was performed, creating several constant speed lines. The results show that the shock structure in the downstream rotor isolates the upstream rotor from the exit conditions until the shock detaches from the leading edge. The shock structure in the upstream rotor then moves, changing the conditions for the downstream rotor. This continues as the compressor is throttled until the shock in the upstream rotor detaches from the leading edge. CFD indicates that this event causes a rapid drop in the mass flow rate, creating a mismatch between stage-one and stage-two that results in compressor instability.Copyright