Improvement of Film Cooling Design for Turbine Vane Leading Edge Considering Combustor Outflow

As the interaction between the combustor and the turbine in the aero-engine continues to increase,the film cooling design considering the combustor swirling outflow has become the research focus.The swirling inflow and high-temperature gas first affect the vane leading edge(LE).However,no practical improved solution for the LE cooling design has been proposed considering the combustor swirling outflow.In this paper,the improved scheme of showerhead cooling is carried out around the two ways of adopting the laid-back-fan-shaped hole and reducing the coolant outflow angle.The film cooling effectiveness(η) and the coolant flow state are obtained by PSP(pressure-sensitive-paint) and numerical simulation methods,respectively.The research results show that the swirling inflow increases the film distribution inhomogeneity by imposing the radial pressure gradient on the vane to make the film excessively gather in some positions.The showerhead film cooling adopts the laid-back-fan-shaped hole to reduce the momentum when the coolant flows out.Although this cooling scheme improves the film attachment and increases the surface-averaged film cooling effectiveness(η_(sur)) by as much as15.4%,the film distribution inhomogeneity increases.After reducing the coolant outlet angle,the wall-tangential velocity of the coolant increases,and the wall-normal velocity decreases.Under the swirl intake condition,both ηand the film distribution uniformity are significantly increased,and the growth of η_(sur) is up to 16.5%.This paper investigates two improved schemes to improve the showerhead cooling under the swirl intake condition to provide a reference for the vane cooling design.

Influence of propagation direction on operation performance of rotating detonation combustor with turbine guide vane

Due to the pressure gain combustion characteristics,the rotating detonation combustor(RDC)can enhance thermodynamic cycle efficiency.Therefore,the performance of gas-turbine engine can be further improved with this combustion technology.In the present study,the RDC operation performance with a turbine guide vane(TGV)is experimentally investigated.Hydrogen and air are used as propellants while hydrogen and air mass flow rate are about 16.1 g/s and 500 g/s and the equivalence ratio is about 1.0.A pre-detonator is used to ignite the mixture.High-frequency dynamic pressure transducers and silicon pressure sensors are employed to measure pressure oscillations and static pressure in the combustion chamber.The experimental results show that the steady propagation of rotating detonation wave(RDW)is observed in the combustion chamber and the mean propagation velocity is above 1650 m/s,reaching over 84%of theoretical Chapman-Jouguet detonation velocity.Clockwise and counterclockwise propagation directions of RDW are obtained.For clockwise propagation direction,the static pressure is about 15%higher in the combustor compared with counterclockwise propagation direction,but the RDW dominant frequency is lower.When the oblique shock wave propagates across the TGV,the pressure oscillations reduces significantly.In addition,as the detonation products flow through the TGV,the static pressure drops up to 32%and 43%for clockwise and counterclockwise propagation process respectively.

Transient simulation of a pump-turbine with misaligned guide vanes during turbine model start-up

Experimental studies of a model pump-turbine S-curve characteristics and its improvement by misaligned guide vanes （MGV） were extended to prototype pump turbine through 3-D transient flow simulations. The unsteady Reynolds-averaged Navier-Stokes equations with the SST turbulence model were used to model the transient flow within the entire flow passage of a reversible pump-turbine with and without misaligned guide vanes during turbine model start-up. The unstable S-curve and its improvement by using misaligned guide vane were verified by model test and simulation. The transient flow calculations were used to clarify the variations of pressure pulse and internal flow behavior in the entire flow passage. The use of misaligned guide vanes can eliminate the S-curve characteristics of a pump-turbine, and can significantly increase the pressure pulse amplitude in the entire flow passage and the runner radial forces during start-up. The MGV only decreased the pulse amplitude on the guide vane suction side when the rotating speed was less than 50% rated speed. The hydraulic reason is that the MGV dramatically changed the flow patterns inside the entire flow passage, and destroyed the symmetry of the flow distribution inside the guide vane and runner.

Multi-scale thermodynamic analysis method for 2D SiC/SiC composite turbine guide vanes

Ceramic Matrix Composite （CMC） turbine guide vanes possess multi-scale stress and strain with inhomogeneity at the microscopic scale. Given that the macroscopic distribution cannot reflect the microscopic stress fluctuation, the macroscopic method fails to meet the requirements of stress and strain analysis of CMC turbine guide vanes. Furthermore, the complete thermodynamic properties of 2D woven SiC/SiC-CMC cannot be obtained through experimentation, Accordingly, a method to calculate the thermodynamic properties of CMC and analyze multi-scale stress and strain of the turbine guide vanes should be established. In this study, the multi-scale thermodynamic analysis is investigated. The thermodynamic properties of Chemical Vapor Infiltration （CVI） pro- cessed SiC/SiC-CMC are predicted by a Representative Volume Element （RVE） model with porosity, leading to the result that the relative error between the calculated in-plane tensile modulus and the experimental value is 4.2%. The macroscopic response of a guide vane under given conditions is predicted. The relative error between the predicted strain on the trailing edge and the experimental value is 9.7%. The calculation of the stress distribution of micro-scale RVE shows that the maximum value of microscopic stress, which is located in the interlayer matrix, is more than 1.5 times that of macroscopic stress in the same direction and the microscopic stress distribution of the interlayer matrix is related to the pore distribution of the composite.

Artificial intelligence aided design of film cooling scheme on turbine guide vane

In recent years,artificial intelligence(AI)technologies have been widely applied in many different fields including in the design,maintenance,and control of aero-engines.The air-cooled turbine vane is one of the most complex components in aero-engine design.Therefore,it is interesting to adopt the existing AI technologies in the design of the cooling passages.Given that the application of AI relies on a large amount of data,the primary task of this paper is to realize massive simulation automation in order to generate data for machine learning.It includes the parameterized three-dimensional(3-D)geometrical modeling,automatic meshing and computational fluid dynamics(CFD)batch automatic simulation of different film cooling structures through customized developments of UG,ICEM and Fluent.It is demonstrated that the trained artificial neural network(ANN)can predict the cooling effectiveness on the external surface of the turbine vane.The results also show that the design of the ANN architecture and the hyper-parameters have an impact on the prediction precision of the trained model.Using this established method,a multi-output model is constructed based on which a simple tool can be developed.It is able to make an instantaneous prediction of the temperature distribution along the vane surface once the arrangements of the film holes are adjusted.

Investigation on Cooling Effectiveness and Aerodynamic Loss of a Turbine Cascade with Film Cooling

This paper describes the numerical study on film cooling effectiveness and aerodynamic loss due to coolant and main stream mixing for a turbine guide vane. The effects of blowing ratio, mainstream Mach number, surface curvature on the cooling effectiveness and mixing loss were studied and discussed. The numerical results show that the distributions of film cooling effectiveness on the suction surface and pressure surface at the same blowing ratio(BR) are different due to local surface curvature and pressure gradient. The aerodynamic loss features for film holes on the pressure surface are also different from film holes on the suction surface.