Endwall aerodynamic losses from turbine components within gas turbine engines

A survey of research on aerodynamic loss investigations for turbine components of gas tuibine engines is presented.Experimental and numerically predicted results are presented from investigations undertaken over the past 65 plus years.Of particular interest are losses from the development of secondary flows from airfoil/endwall interactions.The most important of the airfoilAmdwall secondary flows are passage vortices,counter voitices,and corner vortices.The structure and development of these secondaiy flows are described as they affect aerodynamic perfonnance within and downstream of turbine passage flows in compressible,high speed flows with either subsonic or transonic Mach number distributions,as well as within low-speed,incompressible flows.Also discussed are methods of endwall contouring,and its consequences in regard to airfoil/endwall secondary flows.

Effect of surface roughness on the aerodynamic performance of turbine blade cascade

The effect of surface roughness on the boundary development and loss behavior of turbine blades is investigated with different Reynolds numbers in this paper.The result shows that the velocity profile in boundary layer is plumper on rough surface than on smooth blade.The aerodynamic loss is lowered at low Reynolds number,but becomes significantly large at high Reynolds number.The total pressure loss coefficient of cascade can reach a top increase of 129%for rougher blades comparing with smooth blades at Re=300000.

Aerodynamic and heat transfer performances of a highly loaded transonic turbine rotor with upstream generic rim seal cavity

In turbine disk cavity,rim seals are fitted between the stator and its adjacent rotor disk.A coolant air injected through the turbine disk cavity to prevent the ingress of mainstream hot gases.The purpose of this paper is to investigate numerically the effect of the upstream purge flow on the aero and thermal performances of a high pressure turbine rotor.The investigations are conducted on a generic rim seal cavity inspired from a realistic turbofan engine.Four purge fractions(PF)equal to 0.2%,0.5%,1.0%and 1.5%of the mainstream are considered.The simulations are done by solving the three-dimensional Reynolds averaged Navier-Stokes and energy transport equations.The results include the effect of the PF on the cooling effectiveness,the sealing effectiveness,the secondary flows with losses and the heat transfer behavior,within the cavity and across the rotor passage.The low PF of 0.2%provided a low cooling effectiveness,a moderate sealing effectiveness and minimum losses.The high PF of 1.5%gave a high cooling effectiveness,a best sealing effectiveness and maximum losses.The medium PF of 1.0%supplied a compromise between the aerodynamic and thermal design needs with good cooling and sealing efficiencies and a tolerable level of losses.

Experimental Investigation of Aerodynamic Performance due to Blade Tip Clearance in a Gas Turbine Rotor Cascade

This study examines how the complex flow structure within a gas turbine rotor affects aerodynamic loss. An unshrouded linear turbine cascade was built, and velocity and pressure fields were measured using a 5-hole probe. In order to elucidate the effect of tip clearance, the overall aerodynamic loss was evaluated by varying the tip clearance and examining the total pressure field for each case. The tip clearance was varied from 0% to 4.2% of blade span and the chord length based Reynolds number was fixed at 2×10^(5). For the case without tip clearance, a wake downstream of the blade trailing edge is observed, along with hub and tip passage vortices. These flow structures result in profile loss at the center of the blade span, and passage vortex related losses towards the hub and tip. As the tip clearance increases, a tip leakage vortex is formed, and it becomes stronger and eventually alters the tip passage vortex. Because of the interference of the secondary tip leakage flow with the main flow, the streamwise velocity decreases while the total pressure loss increases significantly by tenfold in the last 30% blade span region towards the tip for the 4.2% tip clearance case. It was additionally observed that the overall aerodynamic loss increases linearly with tip clearance.

Numerical investigation of the interaction between upstream cavity purge flow and main flow in low aspect ratio turbine cascade

In modern gas turbines, rim seal located between the stator-disc and rotor-disc is used to prevent hot-gas ingestion into the inner stage-gap of high pressure turbine. However, the purge flow supplied to the cavity through the rim seal interacts with the main flow, producing additional aerodynamic loss due to the mixing process which plays a significant role in the formation, development and evolution of downstream secondary flow. In this paper, a set of cascade representative of low aspect ratio turbine is selected to numerically investigate the influence of upstream cavity purge flow on the hub secondary flow structure and aerodynamic loss. Cascade with/without upstream cavity and four different purge mass flow rates are all taken into account in this simulation. Then, a deep insight into the loss mechanism of interaction between purge flow and main flow is gained. The results show that the presence of cavity and purge flow has a significant impact on the main flow which not only changes the vortex structure in both the passage and upstream cavity, but also alters the cascade exit flow angle distribution along the spanwise. Moreover, aerodynamic loss in the cascade rises with the increase of purge flow rate while the sealing effect is also enhanced. Therefore, the effect of upstream cavity purge flow must be considered in the process of turbine aerodynamic design. What is more, it is necessary to minimize the purge flow rate in order to reduce aerodynamic loss on the premise of satisfying cooling requirements.

Effects of Width Variation of Pressure-Side Winglet on Tip Flow Structure in a Transonic Rotor

Tip leakage flow has become one of the major triggers for rotating stall in tip region of high loading transonic compressor rotors.Comparing with active flow control method,it’s wise to use blade tip modification to enlarge the stable operating range of rotor.Therefore,three pressure-side winglets with the maximum width of 2.0,2.5 and 3.0 times of the baseline rotor,are designed and surrounded the blade tip of NASA rotor 37,and the three new rotors are named as RPW1,RPW2,and RPW3 respectively.The numerical results show that the width of pressure-side winglet has significant influence on the stall margin and the minimum throttling massflow of rotor,while it produces less effect on the choking massflow and the peak efficiency of new rotors.As the width of the pressure-side winglet increases from new rotor RPW1 to RPW3,the strength of leakage massflow has been attenuated dramatically and a reduction of 20%in leakage massflow rate has appeared in the new rotor RPW3.By contrast,the extended blade tip caused by winglet has not introduced much more aerodynamic losses in tip region of rotor,and the new rotors with different width of pressure-side winglet have the similar peak efficiency to the baseline.The new shape of the leakage channel over blade tip which replaces of the static pressure difference near blade tip has dominated the behavior of the leakage flow in tip gap.As both the new aerodynamic boundary and throat in tip gap have reshaped by the low-velocity flow near the solid wall of extended blade tip,the discharging velocity and massflow rate of leakage flow have been suppressed obviously in new rotors.In addition,the increasing inlet axial velocity at the entrance of new rotor has increased slightly as well,which is attributed to the less blockage in the tip region of new rotor.In consideration of the increased inlet axial velocity and the weakened leakage flow,the new rotor presents an appropriately linear increase of the stall margin when the width of pressure-side winglet increases,and has a nearly 15%increase in new rotor RPW3.