The numerical simulations are used to conduct the comparative study of pin-fins cooling channel and multi-impingement cooling channel on the heat transfer and flow, and to design the multi-impingement channel through the parameters of impinging distance and impingement-jet-plate thickness. The Reynolds number ranges from 1e4 to 6e4. The dimensionless impinging distance is 0.60, 1.68, 2.76, respectively, and the dimensionless impinging-jet-thickness is 0.5, 1.0, 1.5, respectively. The endwall surface, pin-fins surface, impinging-jet-plate surface are the three object surfaces to investigate the channel heat transfer performance. The heat transfer coefficient and augmentation factor are selected to measure the surface heat transfer, and the friction coefficient is chosen to evaluate the channel flow characteristics. The impinging-jet-plate surface owns higher heat transfer coefficient and larger area than pin-fins surface, which are the main reasons to improve the heat transfer performance of multi-impingement cooling channel. Reducing the impinging distance can improve the endwall surface heat transfer obviously and enhance impingement plate surface heat transfer to some extent, decreasing the thickness of impinging-jet-plate can significantly increase its own heat transfer coefficient, which both all increase the cooling air flow loss.
Ceramic Matrix Composites (CMCs) are primary candidates for advanced gas turbine engine application that require intense high temperature tests and validations. Before CMCs used in engine hot sections, a lot of tests need to be done, especially thermal test. A thermal test rig has been set up to simulate the engine turbine thermal environment. Propane gas is used to simulate the practical aviation fuel and compressed air with flow regulator is used as cooling media. The capabilities and limitations of the test facility have been calibrated and discussed in this paper. A CMC turbine vane with internal cooling path was tested on this burning rig. The results showed that the CMC vane could withstand the 1200 ℃ thermal cycling test but the coating was disappeared. It has been proved that such test rig and method could simulate the thermal boundary conditions of turbine vanes and blades.
Control of Mach 1.5 elliptic jet with ventilated triangular tabs is studied experimentally, in the presence of different levels of pressure gradient at the nozzle exit. Three different sets of ventilated tabs with circular, triangular and trapezoidal ventilations were studied. Two tabs were placed, at the ends of major and minor axes, at the exit of the elliptic nozzle of aspect ratio 3.37. The mixing enhancement caused by these tabs was studied in the presence of adverse and favorable pressure gradients, corresponding to nozzle pressure ratio (NPR) from 3 to 8. For Mach 1.5 jet NPR 3 corresponds to 18 % adverse pressure gradient and NPR 8 corresponds to 118 % favorable pressure gradient. The results of ventilated tabs are compared with unventilated truncated triangular tabs of identical geometry. The difference between the mixing promoting efficiency of the unventilated and ventilated tabs is only marginal (around 5–6 %). All tabs cause jet bifurcation and weaken the waves in the jet core. The tab with trapezoidal ventilation, at NPR 3, promotes mixing to an extent of reducing the core to about 92 %. At higher NPRs the mixing caused by unventilated tab is slightly better than the ventilated tabs.
In order to improve compressor performance using a new design method, which originates from the fins on a humpback whale, experimental tests and numerical simulations were undertaken to investigate the influence of the tubercle leading edge on the aerodynamic performance of a linear compressor cascade with a NACA 65–010 airfoil. The results demonstrate that the tubercle leading edge can improve the aerodynamic performance of the cascade in the post-stall region by reducing total pressure loss, with a slight increase in total pressure loss in the pre-stall region. The tubercles on the leading edge of the blades cause the flow to migrate from the peak to the valley on the blade surface around the tubercle leading edge by the butterfly flow. The tubercle leading edge generates the vortices similar to those created by vortex generators, splitting the large-scale separation region into multiple smaller regions.
The following paper presents dynamic leakage rate and coupled interaction for variable speed rotor-labyrinth (LABY) seal, with rotating speed from 18 krpm to 30 krpm. Variable speed rotor vibration characteristics are incorporated into transient computational fluid dynamic (CFD) calculations as boundary conditions of seal flow field to show the real-time effect of rotordynamic in seal flow field. Leakage rate across a variable speed rotor-seal increases with rotor vibration, but this effect is prominent at lower speed than at higher speed. Leakage characteristic is determined by differences in rotor vibration amplitude rather than rotating speed. The results also reveal that aerodynamic forces of labyrinth seal flow field can improve rotor stability, and this interaction between rotor and seal decreases with the increase of rotating speed.
The present numerical investigation of Leading Edge (LE) Nozzle Guide Vane (NGV) is considered with five rows of impingement holes combined with five rows of film cooled for the secondary coolant flow path analysis. The coolant mass flow rate variations in all the LE rows of the film holes externally subjected to the hot main stream were obtained by making a three-dimensional computational analysis of NGV with a staggered array of film cooled rows. The experiments were carried out for the same NGV using Particle Image Velocimetry technique to determine the effused coolant jet exit velocity at the stagnation row of film holes as mentioned in reference [Kukutla PR, Prasad BVSSS. Secondary flow visualization on stagnation row of a combined impingement and film cooled high pressure gas turbine nozzle guide vane using PIV technique, J Visualization, 2017; DOI: 10.1007/s12650-017-0434-6]. In this paper, results are presented for three different mass flow rates ranges from 0.0037 kg/s to 0.0075 kg/s supplied at the Front Impingement Tube (FIT) plenum. And the mainstream velocity 6 m/s was maintained for all the three coolant mass flow rates. The secondary coolant flow distribution was performed from SH1 to SH5 row of film holes. Each row of a showerhead film hole exit coolant mass flow rate varied in proportion to the amount of coolant mass rates supplied at the FIT cooling channel. The corresponding minimum and maximum values and their film hole locations were altered. The same behaviour was continued for the coolant pressure drop and temperature rise from SH1 to SH5 row of film holes. Owing to the interaction between hot main stream and the coolant that effuses out of the film holes, occasional presence of hot gas ingestion was noticed for certain flow rates. This caused nonlinear distribution in mass flow, pressure drop and temperature rise. The minimum flow rate results estimate oxidation of NGV material near the film cooled hole. And the effect of hot gas ingestion on the ejected film cooled jet which would recommends effective oxidation resistant material which in turn leads to better durability of the NGV surface.
Thermal choke is commonly employed in a fixed geometry RBCC combustor to eliminate the need for physically variable exit geometry. This paper proposed detailed numerical studies based on a two-dimensional integration model to characterize thermal choke behaviors driven by various embedded rocket operations in an RBCC engine at Mach 4 in ramjet mode. The influences of different embedded rocket operations as well as the corresponding secondary fuel injection adjustment on thermal choke generation process, the related thermal throat feature, and the engine performance are analyzed. Operations of embedded rocket bring significant effects on the thermal choke behaviors: (1) the thermal throat feature becomes much more irregular influenced by the rocket plume; (2) the occupancy range in the combustor is significantly lengthened; (3) the asynchrony of the flow in different regions accelerating to sonic speed becomes much more significant; (4) as the rocket throttling ratio decreases, the thermal choke position constantly moves upstream integrally, and the heated flow in the top region that is directly affected by the rocket plume reaches sonic speed more rapidly. Finally, we can conclude that appropriate secondary fuel injection adjustment can provide a higher integration thrust for the RBCC engine with the embedded rocket operating, while the thermal choke is stably controlled, and the increased heat release and combustion pressure are well balanced by the variations of pre-combustion shocks in the inlet isolator.
The blade deformation caused by aerodynamic and centrifugal loads during operating makes blade configurations different from their stationary shape. Based on the load incremental approach, a novel pre-deformation method for cold blade shape is provided in order to compensate blade deformation under running. Effect of nonlinear blade stiffness is considered by updating stiffness matrix in response to the variation of blade configuration when calculating deformations. The pre-deformation procedure is iterated till a converged cold blade shape is obtained. The proposed pre-deformation method is applied to a transonic compressor rotor. Effect of load conditions on blade pre-deformation is also analyzed. The results show that the pre-deformation method is easy to implement with fast convergence speed. Neither the aerodynamic load nor centrifugal load can be neglected in blade pre-deformation.