William W. Copenhaver

Three-Dimensional Flow Phenomena in a Transonic, High-Throughflow, Axial-Flow Compressor Stage

A detailed aerodynamic study of a transonic, high-throughflow, single-stage compressor is presented. The compressor stage was comprised of a low-aspect-ratio rotor combined alternately with two different stator designs. Both experimental and numerical studies are conducted to understand the details of the complex flow field present in this stage. Aerodynamic measurements using high-frequency, Kulite pressure transducers and conventional probes are compared with results from a three-dimensional viscous flow analysis. A steady multiple blade row approach is used in the numerical technique to examine the detailed flow structure inside the rotor and the stator passages. The comparisons indicate that many flow field features are correctly captured by viscous flow analysis, and therefore unmeasured phenomena can be studied with some level of confidence.

Stator-Rotor Interactions in a Transonic Compressor— Part 1: Effect of Blade-Row Spacing on Performance

Usually less axial spacing between the blade rows of an axial flow compressor is associated with improved efficiency. However, mass flow rate, pressure ratio, and efficiency all decreased as the axial spacing between the stator and rotor was reduced in a transonic compressor rig. Reductions as great as 3.3% in pressure ratio, and 1.3 points of efficiency were observed as axial spacing between the blade rows was decreased from far apart to close together. The number of blades in the stator blade-row also affected stage performance. Higher stator blade-row solidity led to larger changes in pressure ratio efficiency, and mass flow rate with axial spacing variation. Analysis of the experimental data suggests that the drop in performance is a result of increased loss production due to blade-row interactions. Losses in addition to mixing loss are present when the blade-rows are spaced closer together. The extra losses are associated with the upstream stator wakes and are most significant in the midspan region of the flow.

Stator-Rotor Interactions in a Transonic Compressor: Part 2 — Description of a Loss Producing Mechanism

A previously unidentified loss producing mechanism resulting from the interaction of a transonic rotor blade-row with an upstream stator blade-row is described. This additional loss occurs only when the two blade rows are spaced closer together axially. Time-accurate simulations of the flow and high-response static pressure measurements acquired on the stator blade surface reveal important aspects of the fluid dynamics of the production of this additional loss. At close spacing the rotor bow shock is chopped by the stator trailing edge. The chopped bow shock becomes a pressure wave on the upper surface of the stator that is nearly normal to the flow and that propagates upstream. In the reference frame relative to this pressure wave, the flow is supersonic and thus a moving shock wave that produces an entropy rise and loss is experienced. The effect of this outcome of blade-row interaction is to lower the efficiency, pressure ratio, and mass flow rate observed as blade-row axial spacing is reduced from far to close. The magnitude of loss production is affected by the strength of the bow shock and how much it turns as it interacts with the trailing edge of the stator. At far spacing the rotor bow shock degenerates into a bow wave before it interacts with the stator trailing edge and no significant pressure wave forms on the stator upper surface. For this condition, no additional loss is produced.Copyright

Stator-rotor interactions in a transonic compressor: Part 2: Description of a loss-producing mechanism

A previously unidentified loss producing mechanism resulting from the interaction of a transonic rotor blade row with an upstream stator blade row is described. This additional loss occurs only when the two blade rows are spaced closer together axially. Time-accurate simulations of the flow and high-response static pressure measurements acquired on the stator blade surface reveal important aspects of the fluid dynamics of the production of this additional loss. At close spacing the rotor bow shock is chopped by the stator trailing edge. The chopped bow shock becomes a pressure wave on the upper surface of the stator that is nearly normal to the flow and that propagates upstream. In the reference frame relative to this pressure wave, the flow is supersonic and thus a moving shock wave that produces an entropy rise and loss is experienced. The effect of this outcome of blade-row interaction is to lower the efficiency, pressure ratio, and mass flow rate observed as blade-row axial spacing is reduced from far to close. The magnitude of loss production is affected by the strength of the bow shock and how much it turns as it interacts with the trailing edge of the stator. At far spacing the rotor bow shock degenerates into a bow wave before it interacts with the stator trailing edge and no significant pressure wave forms on the stator upper surface. For this condition, no additional loss is produced.

Unsteady aerodynamic flow phenomena in a transonic compressor stage

A three-dimensional unsteady, viscous aerodynamic analysis has been developed for the e ow inside a transonic, high-through-e ow, single-stage compressor. The compressor stage is comprised of a low-aspectratio rotor and a closely coupled stator. The analysis is based on a numerical method for solving the three-dimensional Navier‐ Stokes equation for unsteady viscous e ow through multiple turbomachinery blade rows. The method solves the fully three-dimensional Navier‐ Stokes equation with an implicit scheme. A two-equation turbulence model with a low Reynolds number modie cation is applied for the turbulence closure. A third-order accurate upwinding scheme is used to approximate convection terms, whereas a second-order-accurate central difference scheme is used for the discretization of the viscous terms. A second-order accurate scheme is employed for the temporal discretization. The numerical method is applied to study the unsteady e owe eld inside a transonic, high-through-e ow, axial compressor stage. The numerical results are compared with available experimental data.

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.

Unsteady Flow and Shock Motion in a Transonic Compressor Rotor

The results of an experimental and numerical comparison of the unsteadiness and shock motion that occurs in the tip region within a modern, low-aspect-ratio, high-through-e ow, axial-e ow transonic fan rotor are presented. The unsteadiness studied here is associated with local phenomena within the blade passage and not related to blade row interaction. Unsteady static pressures were measured at the casing over the rotor that operates at a tip relative Mach number of 1.6. An unsteady three-dimensional Navier ‐ Stokes computational study was performed with tip clearance comparable to the test rotor. The fully three-dimensional, unsteady, Reynolds-averaged Navier ‐ Stokes equations were solved with time steps, each nominally of 2.8 3 10 25 s in duration or approximately e ve times blade pass frequency. The results in the clearance gap were retained from the computational solution and compared with the experimental measurements. Both indicated deterministic unsteadiness near the location of the shock. High levels of unsteadiness were also measured downstream of the shock in the path of the clearance vortex, but this phenomenon was not predicted by the computation. The unsteadiness near the shock was shown to be a result of movement of the shock. The amplitude and frequency of shock position oscillation was estimated from the results to be about 2% chord and 2 kHz, respectively. Analysis of loss caused by the shock unsteadiness suggested that losses because of shock motion were insignie cant relative to the steady shock loss at the relative Mach number studied.

A Three-Dimensional Shock Loss Model Applied to an Aft-Swept, Transonic Compressor Rotor

An analysis of the effectiveness of a three-dimensional shock loss model used in transonic compressor rotor design is presented. The model was used during the design of an all-swept, transonic compressor rotor. The demonstrated performance of the swept rotor, in combination with numerical results, is used to determine the strengths and weaknesses of the model. The numerical results were obtained from a fully three-dimensional Navier-Stokes solver. The shock loss model was developed to account for the benefit gained with three-dimensional shock sweep. Comparisons with the experimental and numerical results demonstrated that shock loss reductions predicted by the model due to the swept shock induced by the swept leading edge of the rotor were exceeded. However, near the tip the loss model underpredicts the loss because the shock geometry assumed by the model remains swept in this region while the numerical results show a more normal shock orientation. The design methods and the demonstrated performance of the swept rotor are also presented. Comparisons are made between the design intent and measured performance parameters. The aft-swept rotor was designed using an inviscid axisymmetric streamline curvature design system utilizing arbitrary airfoil blading geometry. The design goal specific flow rate was 214.7 kg/s/m 2 (43.98 Ibm/sec/ft 2 ), the design pressure ratio goal was 2.042, and the predicted design point efficiency was 94.0. The rotor tip speed was 457.2 m/s (1500 ft/sec). The design flow rate was achieved while the pressure ratio fell short by 0.07. Efficiency was 3 points below prediction, though at a very high 91 percent. At this operating condition the stall margin was 11 percent.

Rotating Stall Performance and Recoverability of a High-Speed 10-Stage Axial Flow Compressor

Design factors which influence rotating stall recoverability of a high-speed multistage compressor are not yet fully understood. A high-speed, 10-stage compressor component was tested while operating in-stall to investigate parameters that affect the overall recoverability of a multistage compressor. The compressor instrumentation and data acquisition procedures were designed to obtain detailed performance data from the compressor while it entered into a rotating stall condition and while it operated in rotating stall. The compressor was tested at different in-stall operating conditions by varying compressor shaft speed, discharge throttle, and variable geometry settings to determine the effect of each variable on rotating-stall performance and recoverability. Test results suggest that the stall cell may not extend the full length of the compressor but instead can be confined to a portion (here the rear stages) of the compressor. When stages are stacked together, as is the case in a multistage compressor, the stalled performance and subsequent recoverability are greatly affected by how well the stages of the compressor are matched. The results also suggest that high-speed flows in the tenth stage may extend in-stall operation causing low recoverability of the overall test compressor at higher shaft speeds.