Reduction Of Total Pressure Loss Research Articles

Sensitivity of an asymmetric 3D diffuser to vortex-generator induced inlet condition perturbations

Modifications of the turbulent separated flow in an asymmetric three-dimensional diffuser due to inlet condition perturbations were investigated using conventional static pressure measurements and velocity data acquired using magnetic resonance velocimetry (MRV). Previous experiments and simulations revealed a strong sensitivity of the diffuser performance to weak secondary flows in the inlet. The present, more detailed experiments were conducted to obtain a better understanding of this sensitivity. Pressure data were acquired in an airflow apparatus at an inlet Reynolds number of 10,000. The diffuser pressure recovery was strongly affected by a pair of longitudinal vortices injected along one wall of the inlet channel using either dielectric barrier discharge plasma actuators or conventional half-delta wing vortex generators. MRV measurements were obtained in a water flow apparatus at matched Reynolds number for two different cases with passive vortex generators. The first case had a pair of counter-rotating longitudinal vortices embedded in the boundary layer near the center of the expanding wall of the diffuser such that the flow on the outsides of the vortices was directed toward the wall. The MRV data showed that the three-dimensional separation bubble initially grew much slower causing a rapid early reduction in the core flow velocity and a consequent reduction of total pressure losses due to turbulent mixing. This produced a 13% increase in the overall pressure recovery. For the second case, the vortices rotated in the opposite sense, and the image vortices pushed them into the corners. This led to a very rapid initial growth of the separation bubble and formation of strong swirl at the diffuser exit. These changes resulted in a 17% reduction in the overall pressure recovery for this case. The results emphasize the extreme sensitivity of 3D separated flows to weak perturbations.

Minimizing Inlet Distortion for Hybrid Wing Body Aircraft

A study of boundary-layer-ingesting flow in an upper-mounted offset inlet in NASA’s hybrid wing body transport concept for achieving environmental and performance requirements has been carried out. This study aims specifically at minimizing flow distortion stemming from the ingested low-momentum fluid to the level acceptable to the operation of fan blades. In this paper, we will focus on using a discrete adjoint method to arrive at an optimized wall geometry, which is parametrically represented by design variables, for transonic turbulent flows described by 3D Navier–Stokes equations supplemented with κ-ω-SST turbulence model. Of special interest herein is the flow physics resulting from optimization, revealing the intricate connections to the remarkable reduction in flow distortion and total pressure losses at the engine face. It is discovered that the counter-rotating vortex pair commonly seen in an S-inlet is eliminated by energizing both the core and boundary layer fluids through the change of wall shape by a series of peaks and valleys. Their exact forms, magnitudes, and locations, unknown a priori, are determined by the discrete adjoint method. The optimal shape, moreover, still holds similar level of superior performance at off-design conditions. The result may suggest a possible paradigm shift in flow control concept, away from disruptive penalty ridden devices, by properly conditioning and guiding the flow.

Separation Control on a Very High Lift Low Pressure Turbine Airfoil Using Pulsed Vortex Generator Jets

Boundary layer separation control has been studied using vortex generator jets (VGJs) on a very high lift, low-pressure turbine airfoil. Experiments were done under high (4%) freestream turbulence conditions on a linear cascade in a low speed wind tunnel. Pressure surveys on the airfoil surface and downstream total pressure loss surveys were documented. Instantaneous velocity profile measurements were acquired in the suction surface boundary layer. Cases were considered at Reynolds numbers (based on the suction surface length and the nominal exit velocity from the cascade) of 25,000 and 50,000. Jet pulsing frequency, duty cycle, and blowing ratio were all varied. Computational results from a large eddy simulation of one case showed reattachment in agreement with the experiment. In cases without flow control, the boundary layer separated and did not reattach. With the VGJs, separation control was possible even at the lowest Reynolds number. Pulsed VGJs were more effective than steady jets. At sufficiently high pulsing frequencies, separation control was possible even with low jet velocities and low duty cycles. At lower frequencies, higher jet velocity was required, particularly at low Reynolds numbers. Effective separation control resulted in an increase in lift and a reduction in total pressure losses. Phase averaged velocity profiles and wavelet spectra of the velocity show the VGJ disturbance causes the boundary layer to reattach, but that it can reseparate between disturbances. When the disturbances occur at high enough frequency, the time available for separation is reduced, and the separation bubble remains closed at all times.

Numerical Study of Effect of Streamwise End Wall Fences on Secondary Flow Losses in Two Dimensional Turbine Rotor Cascade

:In a turbine cascade flow losses due to secondary flow are nearly 33% of total losses. In the present study effect of streamwise end wall fence on secondary flow loss reduction is studied numerically with experimental validation. A fence was fixed normal to the end wall and at half pitch away from the blades. Curvature of the fence was the same as that of the blade camberline. Optimization of the fence height for minimum secondary losses is done numerically using coefficient of secondary kinetic energy as the optimization parameter. A fence of height 1/3rd of inlet boundary layer thickness (δ) was found to be optimum and produces a total pressure loss reduction of 15% compared to cascade without fence. This fence trapped the pressure side leg of horseshoe vortex and reduced maximum overturning near the end wall compared to baseline case. A counter-rotating vortex is produced by this fence which weakens the passage vortex and trailing edge vortex. Climbing of passage vortex on suction surface reduced drastically. Fence of height 1/3rd of inlet boundary layer thickness reduced the three dimensionality of the flow and so the losses.

축류형 송풍기의 익단간극이 성능에 미치는 영향에 관한 연구

Fan performances are obtained with various tip clearance gaps and stagger angles of the rotor. A tested fan is an axial-type fan of which the casing diameter is 806 ㎜. Two different rotors are applied to this test. One is designed on the basis of the free vortex method along the radial direction and the other is designed using the forced vortex method. The operating conditions are varied to the ultimate off-design point as well as the deign point. Overall efficiency, total pressure and input power are compared with the tip clearance gaps and different stagger angle. The experimental results show that changing of the stagger angle has minor influence to the performance when the same rotor is applied. When the tip clearance gap is less than 5% of the rotor span, the overall efficiency, total pressure loss and input power reduction are varied linearly with the variation of the tip clearance gaps. On the design point, the overall efficiency is decreased to the rate of 2.8-2.9 to the increasing of the tip clearance, but the changing rate of the overall efficiency is alleviated when the fan operates at off-design points. In particular, this rate is more quickly declined on a fan with the rotor designed using the forced vortex method. The result of the total pressure shows that the pressure reduction rate is a 0.08-0.1 according to the tip clearance, and additionally the input power reduction rate is a 0.045-0.065 at design point.

Computations of Turbulent Flow and Heat Transfer Through a Three-Dimensional Nonaxisymmetric Blade Passage

The design of a three-dimensional nonaxisymmetric end wall is carried out using three-dimensional numerical simulations. The computations have been conducted both for the flat and contoured end walls. The performance of the end wall is evaluated by comparing the heat transfer and total pressure loss reduction. The contouring is done in such a way to have convex curvature in the pressure side and concave surface in the suction side. The convex surface increases the velocity by reducing the local static pressure, while the concave surface decreases the velocity by increasing the local pressure. The profiling of the end wall is done by combining two curves, one that varies in the streamwise direction, while the other varies in the pitchwise direction. Several contoured end walls are created by varying the streamwise variation while keeping the pitchwise curve constant. The flow near the contoured end wall is seen to be significantly different from that near the flat end wall. The contoured end wall is found to reduce the secondary flow by decreasing radial pressure gradient. The total pressure loss is also lower and the average heat transfer reduces by about 8% compared to the flat end wall. Local reductions in heat transfer are significant (factor of 3). This study demonstrates the potential of three-dimensional end-wall contouring for reducing the thermal loading on the end wall.

Shock Wave/ Boundary-Layer Interaction Control Using Streamwise Slots in Transonic Flows

The effect of streamwise slots on the interaction of a normal shock wave with a turbulent boundary layer has been investigated experimentally at a Mach number of 1.29. The surface-pressure distribution for the controlled interaction was found to feature a distinct plateau. This was caused by a change in shock structure from a typical unseparated normal shock-wave/boundary-layer interaction to a large bifurcated lambda-type shock pattern, which led to a reduction of total pressure losses. A strong spanwise variation of boundary-layer properties was observed downstream of the slots, whereas the modified shock structure was relatively two-dimensional. Surface flow visualization confirmed that the slots introduced a region of recirculation into the boundary layer, similar to passive control with uniform surface ventilation. Surface flow visualization revealed the presence of a pair of counter-rotating vortices, confirmed by crossflow velocity measurements. Because of the reduction of total pressure losses, streamwise slots can reduce aircraft wave drag at transonic cruise while incurring only small viscous penalties. A similar control device can also be of use in supersonic intakes where total pressure losses limit engine performance

A study on an axial-type 2-D turbine blade shape for reducing the blade profile loss

Losses on the turbine consist of the mechanical loss, tip clearance loss, secondary flow loss and blade profile loss etc.,. More than 60 % of total losses on the turbine is generated by the two latter loss mechanisms. These losses are directly related with the reduction of turbine efficiency. In order to provide a new design methodology for reducing losses and increasing turbine efficiency, a two-dimensional axial-type turbine blade shape is modified by the optimization process with two-dimensional compressible flow analysis codes, which are validated by the experimental results on the VKI turbine blade. A turbine blade profile is selected at the mean radius of turbine rotor using on a heavy duty gas turbine, and optimized at the operating condition. Shape parameters, which are employed to change the blade shape, are applied as design variables in the optimization process. Aerodynamic, mechanical and geometric constraints are imposed to ensure that the optimized profile meets all engineering restrict conditions. The objective function is the pitchwise area averaged total pressure at the 30 % axial chord downstream from the trailing edge. 13 design variables are chosen for blade shape modification. A 10.8 % reduction of total pressure loss on the turbine rotor is achieved by this process, which is same as a more than 1 % total-to-total efficiency increase. The computed results are compared with those using 11 design variables, and show that optimized results depend heavily on the accuracy of blade design.

The interaction of shock waves and heat addition in the design of supersonic combustors

Analytical models are needed that will permit the prediction of thrust efficiency of a supersonic combustor as a function of combustor geometry and heat release. An earlier, more general procedure for the one-dimensional analysis of supersonic combustors was developed on the basis of exponential pressure-area dependence. This paper extends that analysis by establishing three regimes for specific pseudo-one-dimensional solutions for thrust efficiency: (1) a low-heat-release regime in which the flow can be considered essentially shock-free; (2) an intermediate regime in which an oblique shock wave is sustained with a pressure ratio equal to or greater than that required for turbulent boundary-layer separation; and (3) a high-heat-release regime in which the combustion is preceded by a normal shock. The key to solution is the formulation for the wall-pressure force needed for the momentum equation. With it, a set of integral equations can be solved simultaneously for relationships between pressure ratio, total temperature ratio and area ratio across the combustor in the three regimes. It is argued that the most probable solutions are those that yield the lowest back pressure without violating the entropy limit (second law of thermodynamics). It is shown that such limiting processes, including shock waves, can yield higher over-all efficiencies than shock-free solutions, because the total pressure loss across the shock is more than compensated by the reduction in total pressure loss due to heat addition, which occurs at a lower Mach number. Results of a typical substantiating experiment falling in this class are presented, for a combustor-inlet Mach number of 3.2 and an alkyl borane fuel.