Simulation of three-dimensional viscous flow in turbomachinery geometries using a solution-adaptive unstructured mesh methodology

This paper presents a numerical method for the simulation of flow in turbomachinery blade rows using a solution-adaptive mesh methodology. The fully three-dimensional, compressible, Reynolds-averaged Navier-Stokes equations with k-ε turbulence modeling (and low Reynolds number damping terms) are solved on an unstructured mesh formed from tetrahedral finite volumes. At stages in the solution, mesh refinement is carried out based on flagging cell faces with either a fractional variation of a chosen variable (like Mach number) greater than a given threshold or with a mean value of the chosen variable within a given range. Several solutions are presented, including that for the highly three-dimensional flow associated with the corner stall and secondary flow in a transonic compressor cascade, to demonstrate the potential of the new method.

The development of turbine exit flow in a swan-necked inter-stage diffuser

This paper describes both the migration and dissipation of flow phenomena downstream of a transonic high-pressure turbine stage. The geometry of the HP stage exit duct considered is a swan-necked diffuser similar to those likely to be used in future engine designs. The paper contains results both from an experimental programme in a turbine test facility and from numerical predictions. Experimental data was acquired using three fast-response aerodynamic probes capable of measuring Mach number, whirl angle, pitch angle, total pressure and static pressure. The probes were used to make time-resolved area traverses at two axial locations downstream of the rotor trailing edge. A 3D time-unsteady viscous Navier-Stokes solver was used for the numerical predictions. The unsteady exit flow from a turbine stage is formed from rotordependent phenomena (such as the rotor wake, the rotor trailing edge recompression shock, the tip-leakage flow and the hub secondary flow) and vane-rotor interaction dependant phenomena. This paper describes the time-resolved behaviour and three-dimensional migration paths of both of these phenomena as they convect downstream. It is shown that the inlet flow to a downstream vane is dominated by two corotating vortices, the first caused by the rotor tip-leakage flow and the second by the rotor hub secondary flow. At the inlet plane of the downstream vane the wake is extremely weak and the radial pressure gradient is shown to have caused the majority of the high loss wake fluid to be located between the mid-height of the passage and the casing wall. The structure of the flow indicates that between a high pressure stage and a downstream vane simple two-dimensional blade row interaction does not occur. The results presented in this paper indicate that the presence of an upstream stage is likely to significantly alter the structure of the secondary flow within a downstream vane. The paper also shows that vane-rotor interaction within the upstream stage causes a 10° circumferential variation in the inlet flow angle of the 2nd stage vane.

The development of aerodynamic performance measurements in a transient test facility

Transient test facilities offer the potential for the simultaneous study of turbine aerodynamic performance, unsteady flow phenomena and the heat transfer characteristics of a turbine stage. This paper describes the development of aerodynamic performance measurement techniques in the Oxford Rotor Facility (ORF). The solutions to the technological issues involved with transient testing presented in this paper are expected to achieve levels of precision uncertainty comparable with traditional steady flow test rigs. The theoretical background to the measurement of aerodynamic performance is presented together with a comprehensive pre-test uncertainty analysis. The instrumentation scheme for the measurement of stage mass flow rate is discussed in detail, the measurements of shaft power, total inlet enthalpy, and stage pressure ratio are also outlined. The current working section features a 62% scale, 1-1/2 stage, high-pressure shroudless transonic turbine. The required inlet flow conditions are provided by an Isentropic Light Piston Tunnel (ILPT) with a quasi-steady state run time of approximately 70ms. The testing is conducted at engine representative specific speed, pressure ratio, gas-to-wall temperature ratio, Mach number and Reynolds number.

The effects of a trip wire and unsteadiness on a high speed highly loaded low-pressure turbine blade

This paper presents the effect of a single spanwise 2D wire upon the downstream position of boundary layer transition under steady and unsteady inflow conditions. The study is carried out on a high turning, high-speed, low pressure turbine (LPT) profile designed to take account of the unsteady flow conditions. The experiments were carried out in a transonic cascade wind tunnel to which a rotating bar system had been added. The range of Reynolds and Mach numbers studied includes realistic LPT engine conditions and extends up to the transonic regime. Losses are measured to quantify the influence of the roughness with and without wake passing. Time resolved measurements such as hot wire boundary layer surveys and surface unsteady pressure are used to explain the state of the boundary layer. The results suggest that the effect of roughness on boundary layer transition is a stability governed phenomena, even at high Mach numbers. The combination of the effect of the roughness elements with the inviscid Kelvin-Helmholtz instability responsible for the rolling up of the separated shear layer (Stieger [1]) is also examined. Wake traverses using pneumatic probes downstream of the cascade reveal that the use of roughness elements reduces the profile losses up to exit Mach numbers of 0.8. This occurs with both steady and unsteady inflow conditions.

Evaluation of wave drag reduction by flow control

An analytical expression is proposed to estimate the wave drag of an aerofoil equipped with shock control. The analysis extends the conventional approach for a single normal shock wave, based on the knowledge that all types of successful shock control on transonic aerofoils cause bifurcated λ-shock structures. The influence of surface curvature on the λ-shock structure has been taken into account. The extended method has been found to produce fairly good agreement with the results obtained by CFD methods while requiring negligible computational effort. This new formulation is expected to be beneficial in the industrial design process of transonic aerofoils and wings where a large number of computational simulations have to be performed.

Unsteady shock motion on a NACA0012 aerofoil at low reduced frequencies

Using transonic blowdown windtunnel experiments, the 2D unsteady shock motion on a NACA0012 aerofoil is examined at various frequencies typical for helicopter blades in forward flight. The aerofoil is subjected to freestream velocities oscillating periodically between M = 0.66 and M = 0.77. Unsteady pressure traces and schlieren images are analyzed over a range of low reduced frequencies to provide information on shock location and strength throughout the cycle. Unsteady effects were noticeable even at very low reduced frequencies (down to O(0.01). However, through the range of frequencies investigated, and within experimental error, the unsteady shock location showed no discernible lag compared to the quasi-steady behaviour. On the other hand, significant variations were observed in shock strengths with the upstream running part of the cycle (decreasing Mach number) displaying considerably stronger shocks than during the accelerating part of the cycle. It could be shown that this variation in shock strength is primarily caused by the shock motion modifying the relative shock Mach number. As a result is was possible to use the quasi-steady results to predict the unsteady shock behaviour at the frequencies investigated here (below 0(0.1)).

Unsteady shock motion on a NACA0012 aerofoil at low reduced frequencies

Using transonic blowdown windtunnel experiments, the 2D unsteady shock motion on a NACA0012 aerofoil is examined at various frequencies typical for helicopter blades in forward flight. The aerofoil is subjected to freestream velocities oscillating periodically between M = 0.66 and M = 0.77. Unsteady pressure traces and schlieren images are analyzed over a range of low reduced frequencies to provide information on shock location and strength throughout the cycle. Unsteady effects were noticeable even at very low reduced frequencies (down to O(0.01). However, through the range of frequencies investigated, and within experimental error, the unsteady shock location showed no discernible lag compared to the quasi-steady behaviour. On the other hand, significant variations were observed in shock strengths with the upstream running part of the cycle (decreasing Mach number) displaying considerably stronger shocks than during the accelerating part of the cycle. It could be shown that this variation in shock strength is primarily caused by the shock motion modifying the relative shock Mach number. As a result is was possible to use the quasi-steady results to predict the unsteady shock behaviour at the frequencies investigated here (below 0(0.1)).

Laminar to turbulent flow transition measurements using an array of SOI-CMOS MEMS wall shear stress sensors

The successful utilization of an array of silicon on insulator complementary metal oxide semiconductor (SOICMOS) micro thermal shear stress sensors for flow measurements at macro-scale is demonstrated. The sensors use CMOS aluminum metallization as the sensing material and are embedded in low thermal conductivity silicon oxide membranes. They have been fabricated using a commercial 1 μm SOI-CMOS process and a post-CMOS DRIE back etch. The sensors with two different sizes were evaluated. The small sensors (18.5 ×18.5 μm2 sensing area on 266 × 266 μm2 oxide membrane) have an ultra low power (100 °C temperature rise at 6mW) and a small time constant of only 5.46 μs which corresponds to a cut-off frequency of 122 kHz. The large sensors (130 × 130 μm2 sensing area on 500 × 500 μm2 membrane) have a time constant of 9.82 μs (cut-off frequency of 67.9 kHz). The sensors' performance has proven to be robust under transonic and supersonic flow conditions. Also, they have successfully identified laminar, separated, transitional and turbulent boundary layers in a low speed flow. © 2008 IEEE.

Experimental study into the flow physics of three-dimensional shock control bumps

This paper describes a fundamental experimental study of the flow structure around a single three-dimensional (3D) transonic shock control bump (SCB) mounted on a flat surface in a wind tunnel. Tests have been carried out with a Mach 1.3 normal shock wave located at a number of streamwise positions relative to the SCB. Details of the flow have been studied using the experimental techniques of schlieren photography, surface oil flow visualization, pressure sensitive paint, and laser Doppler anemometry. The results of the work build on the findings of previous researchers and shed new light on the flow physics of 3D SCBs. It is found that spanwise pressure gradients across the SCB ramp and the shape of the SCB sides affect the magnitude and uniformity of flow turning generated by the bump, which can impact on the spanwise propagation of the quasi-two-dimensional (2D) shock structure produced by a 3DSCB. At the bump crest, vortices can form if the pressure on the crest is significantly lower than at either side of the bump. The trajectories of these vortices, which are relatively weak, are strongly influenced by any spanwise pressure gradients across the bump tail. Asignificant difference between 2D and 3D SCBs highlighted by the study is the impact of spanwise pressure gradients on 3D SCB performance. The magnitude of these spanwise pressure gradients is determined largely by SCB geometry and shock position. Copyright © 2011 by the American Institute of Aeronautics and Astronautics, Inc.

Shock control bump robustness enhancement

Robustness enhancement for Shock Control Bumps (SCBs) on transonic wings is an ongoing topic because most designs provide drag savings only in a relatively small band of the airfoil polar. In this paper, different bump shapes are examined with CFD methods which are validated first by comparison with wind tunnel results. An evaluation method is introduced allowing the robustness assessment of a certain design with little computational effort. Shape optimizations are performed to trim SCB designs to maximum performance on the one hand and maximum robustness on the other hand. The results are analysed and different and parameters influencing the robustness are suggested. Copyright © 2012 by Klemens Nuebler.

Toward designing with three-dimensional bumps for lift/drag improvement and buffet alleviation

The desire to design more efficient transport aircraft has led to many different attempts to minimize drag. One approach is the use of three-dimensional shock control bumps, which have gained popularity in the research community as simple, efficient and robust devices capable of reducing the wave drag of transonic wings. This paper presents a computational study of the performance of three-dimensional bumps, relating key bump design variables to the overall wing aerodynamic performance. An efficient parameterization scheme allows three-dimensional bumps to be directly compared to two-dimensional designs, indicating that two-dimensional bumps are capable of greater design point aerodynamic performance in the transonic regime. An advantage of three-dimensional bumps lies in the production of streamwise vortices, such that, while two-dimensional bumps are capable of superior performance near the design point, three-dimensional bumps are capable of breakingup regions of separated flow at high Mach numbers, suggesting improvement in terms of buffet margin. A range of bump designs are developed that exhibit a tradeoff between design point aerodynamic efficiency and improvementinbuffet margin, indicating the potential for bespoke designs to be generated for different sections of a wing based on its flow characteristics. Copyright © 2012 by Jeremy Eastwood and Jerome Jarrett.