Rotor Shock Research Articles

Axial Compressor Transonic Rotor Shock Interaction With Upstream Stator Row

Abstract Modern multistage axial compressors are designed for high aerodynamic loading and wide range with front stages rotor inlet relative Mach number above one over a large portion of span. Rotor shock waves propagate and reflect in the upstream stator where they provoke extra aerodynamic losses and aeromechanic risks. This paper discusses stator–rotor shock interactions using both steady and unsteady numerical simulations. The computations focus on three different stage designs with the same capacity and total pressure ratio, but different axial gaps and load distribution in stator and rotor. Speedlines from steady and unsteady computations compare the designs, and the detailed flow fields are analyzed at design and close to stall conditions. The steady and unsteady calculations predict similar stage performance, but different stator–rotor loss split. The analyses reveal how the rotor shocks orientation with respect to the upstream stator camber line controls the intensity of shocks interaction with the stator row and its propagation upstream. Finally, the paper indicates the ways to reduce the shock-driven rotor–stator interaction and suggests simple criteria to determine when to switch from steady to unsteady calculations for a reliable transonic stage performance prediction.

Variable pitch fan aerodynamic design for reverse thrust operation

Variable pitch low pressure ratio fans could enable higher propulsive efficiency and eliminate the need for heavy thrust reversers. In this paper, the effects of the fan rotor design on reverse thrust capability have been explored by varying key parameters of NASA’s Advanced Ducted Propulsor (ADP) while maintaining cruise point performance. Reverse thrust performance has been assessed using RANS single passage CFD of the variable pitch fan system with an extended domain. This computational approach has been validated using NASA Stage 51B, an experimental variable pitch fan test case. Predicted total pressure and total pressure ratios for this case were found to agree with test data within experimental error, except where large tip region separations occurred at high incidence. Applying individual variations to rotor pitch-to-chord, radial loading distribution, and chordwise camber distribution generated changes to rotor incidence, blockage, and peak surface Mach numbers in reverse operation. An increase in gross reverse thrust of up to 8% was achieved through reductions in rotor pitch-to-chord due to improved loading and lower shock Mach numbers. Controlling section camber distributions was used to move the rotor shock downstream and was found to increase reverse thrust by up to 3%. Increasing rotor loading at the mid span relative to the tip resulted in high tip incidence and a 50% reduction in gross reverse thrust across all rotor speeds. This suggests that tip loaded designs are required for high levels of reverse thrust.

Design and flow rate improvement of a small aeroengine turbine

The need of small high thrust-to-weight ratio aeroengine is growing. One way to meet this is higher through flow rate of engine. Higher engine mass flow means higher turbine mass flow. In the paper, a small turbine is designed and then promoted with higher flow rate. First the definition of universal turbine through flow capability is discussed and formulated. Then the design and flow rate promotion process of a turbine is introduced. And three design results are obtained and compared, marked MI, MII, and MIII. Among them, MI is original design result with highest efficiency of 0.8899 but doesn’t meet mass flow requirement. MII and MIII are two retrofitted turbines based on MI both with improved mass flow rate by about 3.40%. Different flow rate improvement methods are implemented for the two turbines. S shape endwall is applied to MII stator shroud and MIII turns down stator blade exit angle a bit. MII has higher efficiency than MIII but through flow capability is in fact lowered. Efficiency decline has been seen in both two redesigned turbines compared to MI. The reason possibly lies in the increased rotor shock loss found by detailed CFD flow analysis. Finally, MII is regarded as the best result of which efficiency is 0.8884.

Analysis and Prediction of Shock-Induced Vortex Circulation in Transonic Compressors

Multiple high-fidelity time-accurate computational fluid dynamics simulations were performed to investigate the effects of upstream stator loading and rotor shock strength on vortex shedding characteristics in a single-stage transonic compressor. Three loadings on the upstream stator row of decreased, nominal, and increased loading in conjunction with three axial spacings of close, mid, and far were studied for this analysis. The time-accurate urans code turbo was used to generate periodic, quarter annulus simulations of the blade row interaction (BRI) compressor rig. It was observed that vortex shedding was synchronized to the passing of a rotor bow shock. Results show that vortex strength increases linearly with stator loading and rotor bow shock strength. “Normal” and “large” shock-induced vortices formed on the stator trailing edge (TE) immediately after the shock passing, but the large vortices were strengthened at the TE due to a low-velocity region on the suction surface. This low-velocity region was generated upstream on the suction surface from a shock-induced thickening of the boundary layer or separation bubble. The circulation of the large vortices was greater than the normal vortices by a factor of 1.7, 1.83, and 2.04 for decreased, nominal, and increased deswirler loadings. At decreased loading, only 24% of the measured vortices were considered large, while at nominal loading 58% were large. A model was developed to predict shock-induced vortex circulation from a known rotor bow shock strength, stator diffusion factor, and near-wake parameters. The model predicts the average vortex circulation very well, with 5% difference between predicted and measured values. An understanding of the unsteady interactions associated with blade loading and rotor shock strength in transonic stages will help compressor designers account for unsteady flow physics at design and off-design operating conditions.

Effects of Boundary-Layer Ingestion on the Aero-Acoustics of Transonic Fan Rotors

The use of boundary-layer-ingesting, embedded propulsion systems can result in inlet flow distortions where the interaction of the boundary-layer vorticity and the inlet lip causes horseshoe vortex formation and the ingestion of streamwise vortices into the inlet. A previously-developed body-force-based fan modeling approach was used to assess the change in fan rotor shock noise generation and propagation in a boundary-layer-ingesting, serpentine inlet. This approach is employed here in a parametric study to assess the effects of inlet geometry parameters (offset-to-diameter ratio and downstream-to-upstream area ratio) on flow distortion and rotor shock noise. Mechanisms related to the vortical inlet structures were found to govern changes in the rotor shock noise generation and propagation. The vortex whose circulation is in the opposite direction to the fan rotation (counter-swirling vortex) increases incidence angles on the fan blades near the tip, enhancing noise generation. The vortex with circulation in the direction of fan rotation (co-swirling vortex) creates a region of subsonic relative flow near the blade tip radius that decreases the sound power propagated to the far-field. The parametric study revealed that the overall sound power level at the fan leading edge is set by the ingested streamwise circulation, and that for inlet designs in which the streamwise vortices are displaced away from the duct wall, the sound power at the upstream inlet plane increased by as much as 9 dB. By comparing the far-field noise results obtained to those for a conventional inlet, it is deduced that the changes in rotor shock noise are predominantly due to the ingestion of streamwise vorticity.

Differential evolution based soft optimization to attenuate vane–rotor shock interaction in high-pressure turbines

This article presents a soft computing methodology to design turbomachinery components experiencing strong shock interactions. The study targets a reduction of unsteady phenomena using evolutionary optimization with robust, high fidelity, and low computational cost evaluations. A differential evolution (DE) algorithm is applied to optimize the transonic vane of a high-pressure turbine. The vane design candidates are examined by a cost-effective Reynolds-averaged Navier–Stokes (RANS) solver, computing the downstream pressure distortion and aerodynamic efficiency. A reduction up to 55% of the strength of the shock waves propagating downstream of the stand-alone vane was obtained. Subsequently to the vane optimization, unsteady computations of the vane–rotor interaction were performed using a non-linear harmonic (NLH) method. Attenuation above 60% of the unsteady forcing on the rotor (downstream of the optimal vane) was observed, with no stage-efficiency abatement. These results show the effectiveness of the proposed soft optimization to improve unsteady performance in modern turbomachinery exposed to strong shock interactions.

A Surrogate-Based Approach to Reduced-Order Dynamic Stall Modeling

The surrogate-based recurrence framework (SBRF) approach to reduced-order dynamic stall modeling associated with pitching/plunging airfoils subject to fixed or time-varying freestream Mach numbers is described. The SBRF is shown to effectively mimic full-order two-dimensional computational fluid dynamics solutions for unsteady lift, moment, and drag, but at a fraction of the computational cost. In addition to accounting for realistic helicopter rotor blade dynamics, it is shown that the SBRF can model advancing rotor shock induced separation as well as retreating blade stall associated with excessive angles of attack. Therefore, the SBRF is ideally suited for a variety of rotary-wing aeroelasticity and active/passive design optimization studies that require high-fidelity aerodynamic response solutions with minimal computational expense.

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. [S0889-504X(00)00303-2]

Investigation of Flow Phenomena in a Transonic Fan Rotor Using Laser Anemometry

Several flow phenomena, including flow field periodicity, rotor shock oscillation, and rotor shock system geometry have been investigated in a transonic low aspect ratio fan rotor using laser anemometry. Flow periodicity is found to increase with increasing rotor pressure rise and to correlate with blade geometry variations. Analysis of time-accurate laser anemometer data indicates that the rotor shock oscillates about its mean location with an amplitude of 3–4 percent of rotor chord. The shock surface is nearly two-dimensional for levels of rotor pressure rise at and above the peak efficiency level but becomes more complex for lower levels of pressure rise. Spanwise shock lean generates radial flows due to streamline deflection in the hub-to-shroud streamsurface.

Unsteady Losses in Transonic Compressors

A newly developed probe has been used to measure the time-resolved total temperature and pressure in a transonic compressor. The investigation revealed the presence of large fluctuations in the blade-to-blade core flow occurring at a frequency of three to four times blade passing. These variations are not steady in the rotor relative frame and are thought to be caused by unsteadiness in the rotor shock system, driven perhaps by a shed vortex in the blade wake. The loss associated with this unsteadiness is calculated and the implications for stage performance are discussed.

Laser Anemometer Measurements in a Transonic Axial Flow Compressor Rotor

A laser anemometer system employing an efficient data acquisition technique has been used to make measurements upstream, within, and downstream of the compressor rotor. A fluorescent dye technique allowed measurements within endwall boundary layers. Adjustable laser beam orientation minimized shadowed regions and enabled radial velocity measurements outside of the blade row. The flow phenomena investigated include flow variations from passage to passage, the rotor shock system, three-dimensional flows in the blade wake, and the development of the outer endwall boundary layer. Laser anemometer measurements are compared to a numerical solution of the streamfunction equations and to measurements made with conventional instrumentation.

Development of Sonic Inlets for Turbofan Engines

Results of a current program for sonic inlet technology development are presented. This program includes configuration and mechanical design selection of concepts, aerodynamic design description of the models, and results of test evaluation. In the test program several sonic inlet concepts were tested and compared for aerodynamic and acoustic performance. Results of these comparative evaluations are presented. Near-field measurements were taken inside the sonic inlet on several inlet models. Results of these tests are discussed with respect to the effect of Mach number gradients on noise attenuation and rotor shock wave attenuation, and boundary-layer effects on noise propagation. The test facilities and experimental techniques employed are described briefly.