Design of Duct Passages for an Air Turbine Starter Test Rig

In a typical air turbine starter (ATS) engine testing application, compressed air is supplied to the turbine by means of an inlet duct usually with a 90 degree bend and discharged from the turbine into the exhaust chimney through a combination of two duct passages. The primary duct is integral to the engine for connecting to the containment ring. The secondary duct is a transition piece for connecting to the exhaust chimney. As these ducts consume additional pressure and adversely affect the performance of the ATS under test. The design of pressure-efficient outlet ducts is therefore essential, and is the topic of present study. The aerodynamic performance of the overall passage depends on the (i) angle of bend, (ii) the shape of the connecting bolt, (iii) the outlet area and shape of the exhaust duct transiting between the bend and the chimney. Combinations of different angular bends, different shaped bolts and varying size of transition pieces are analyzed using the enterprise version of CFD tool, ANSYS. Three dimensional mesh independent simulations using k-epsilon turbulence model are carried out for a combined geometry of inlet duct, rotor-stator combination, outlet ducts together with the bolts. A combination of the duct passages that has resulted in lowest possible pressure drop is suggested as result of the study i.e. the 90 degree bend duct gives 9% pressure difference between inlet and outlet and this might slightly affect the efficiency of the air turbine stator, however the mass flow rate values remains similar to the stator inlet mass flow rate. Hence the 90 degree bend duct is suitable for the test rig. The static pressure loss and total pressure gain is about 0.04% and −0.004% respectively for baseline and aggressive duct of stator and rotor, hence the baseline duct profile is better than aggressive duct. Among different shapes of connecting bolt, the baseline geometry gives slightly lower efficiency of 85.6% when compared to all other models. But due to manufacturing feasibility the baseline geometry is preferred. Exhaust duct model 7 gives pressure drop as 0.062 bar twice the amount of pressure drop in model 6, but it does not affect the efficiency of air turbine starter. The shapes and sizes of the bend, bolts and the transition piece are recommended.

Investigation of the Flow In Cold Condition at the Exit of a Supersonic Combustor Test Bench

For studies of hypersonic flows and supersonic combustion in ground test facilities, three devices can be used as ram accelerators, shock tunnels and supersonic combustor test benches. These devices can reproduce, on the ground, similar conditions to those in real flight at a certain altitude and speed. In the case of the supersonic combustor test bench (SCTB), it simulates the same flow conditions inside the combustor of a scramjet. The SCTB consists basically of a combustion chamber or vitiated air generator unit, where the air is heated, and a nozzle, where the air is accelerated to the desired test speed. The supersonic combustor to be tested is directly coupled to the nozzle exit of the SCTB. Ultimately, it was necessary to use a transition piece to connect the nozzle to the combustor to be tested, because the nozzle exit has a circular section and the combustor entrance has a rectangular one. This work aims to present the process of characterizing the cold flow along the SCTB of the Institute for Advanced Studies (IEAv) using the schlieren technique. The interference of the transition piece in obtaining the required flow conditions at the exit of the SCTB or the entrance of the combustor was mainly evaluated.

Impact of a Combustor–Turbine Interface on Turbine Vane Aerodynamics and Film Cooling

Optimizing the aerothermal performance of the combustor–turbine interface is an important factor in enhancing the efficiency of heavy-duty gas turbines. Also, it is a key requirement to fulfill the lifetime in this hottest area of the gas turbine. Typically transition pieces of can combustors induce a highly nonuniform swirling flow at the turbine inlet. In order to better understand the impact of the nonuniform combustor flow at the first stage vanes, a combined experimental and numerical study was carried out. The experimental facility consisted of a high-speed linear cascade with four vane passages, including an upstream transition piece, which was representative of a heavy-duty gas turbine can combustor–turbine interface geometry. The experiments were conducted at engine representative Mach numbers, and film cooling effectiveness measurements were performed at three different blowing ratios. Computational fluid dynamics (CFD) Reynolds-averaged Navier–Stokes simulations were undertaken using a commercial flow solver. The numerical model was first validated with the experimental data, using inlet traverse five-hole probe measurements, pressure taps along the airfoil perimeter, and oil flow visualization results. The investigation shows that the position of the vane relative to the combustor transition piece has a significant impact on the vane aerodynamics and also film cooling behavior. This understanding was important for a robust first vane aerothermal design of the GT36.

Numerical Analysis of Stress Distribution in Offshore Wind Turbine M72 Bolted Connections

The use of bolted joints to connect the transition piece and monopile is nowadays widely applied in the offshore wind industry. Traditionally, grouted connections were used in the early generation of offshore wind turbines, but the experienced failures in such connections led to an increased tendency towards bolted flange connections to join the transition piece and monopile in the new generation of offshore wind turbines. The bolts used for this purpose have high strength and huge sizes, and are subjected to a preload that is applied during the tightening process. The present study is focused on the analysis of preload effects on stress distribution in M72 bolted connections by considering different friction coefficients between the bolt and nut threads. The bolt is considered to be made of grade 10.9 steel, whereas the nut is assumed to be made of grade 8.8 steel, which is a softer material. Using the finite element commercial software package Abaqus, numerical models were developed and analysed to establish trends for stress distribution and plastic strains during the bolt tightening process, and to quantify stress concentration factors in individual engaged threads.

Investigation of the Effect of Perforated Sheath on Thermal-Flow Characteristics Over a Gas Turbine Reverse-Flow Combustor—Part 1: Experiment

Reverse-flow combustors have been used in heavy, land-based gas turbines for many decades. A sheath is typically installed over the external walls of the combustor and transition piece to provide enhanced cooling through hundreds of small impinging cooling jets, followed by a strong forced convection channel flow. However, this cooling is at the expense of a large pressure loss. With the modern advancements in metallurgy and thermal-barrier coating technologies, it may become possible to remove this sheath to recover the pressure loss without causing thermal damage to the combustor chamber and the transition piece walls. However, without the sheath, the flow inside the dump diffuser may exert nonuniformly reduced cooling on the combustion chamber and transition piece walls. The objective of this paper is to investigate the difference in flow pattern, pressure drop, and heat transfer distribution in the dump diffuser and over the outer surface of the combustor with and without a sheath. Both experimental and computational studies are performed and presented in Part 1 and Part 2, respectively. The experiments are conducted under low pressure and temperature laboratory conditions to provide a database to validate the computational model, which is then used to simulate the thermal-flow field surrounding the combustor and transition piece under elevated gas turbine operating conditions. The experimental results show that the pressure loss between the dump diffuser inlet and exit is 1.15% of the total inlet pressure for the non-sheathed case and 1.9% for the sheathed case. This gives a 0.75 percentage point (or 40%) reduction in pressure losses. When the sheath is removed in the laboratory, the maximum increase of surface temperature is about 35%, with an average increase of 13–22% based on the temperature scale of 23 K, which is the difference between the bulk inlet and the outlet temperatures.

Investigation of the Effect of Perforated Sheath on Thermal-Flow Characteristics Over a Gas Turbine Reverse-Flow Combustor—Part 2: Computational Analysis

The objective of Part 2 is to employ a computational scheme to investigate the difference in flow pattern, pressure drop, and heat transfer in a gas turbine’s dump diffuser and over the outer surface of the combustor with and without a sheath. Both experimental and computational studies are performed. In Part 1, the experiments were conducted under low pressure and temperature laboratory conditions to provide a database to validate the computational model, which was then used to simulate the thermal-flow field surrounding the combustor and transition piece under elevated gas turbine operating conditions. For laboratory conditions, the computational fluid dynamics (CFD) results show that (a) the predicted local static pressure values are higher than the experimental data but the prediction of the global total pressure losses matches the experimental data very well; (b) the total pressure losses are within 3% of the experimental values; and (c) the temperature difference between the sheathed and non-sheathed cases is in the range of 5–10 K or 16–32% based on the temperature scale between the highest and lowest temperatures in the computational domain. Under the elevated pressure and temperature conditions in real gas turbine, removing the sheath can achieve a significant pressure recovery of approximately 3% of the total pressure, but it will be subject to a wall temperature increase of about 500 K (900 °F or a 36% increase) on the outer radial part of the transition piece, where the flow is slow due to diffusion and recirculation in the large dump diffuser cavity near the turbine end. If modern advanced materials or coatings could sustain a wall temperature of about 500 K higher than those currently available, the sheath could be removed. Otherwise, removal of the sheath is not recommended.

Analysis of Tightening Sequence Effects on Preload Behaviour of Offshore Wind Turbine M72 Bolted Connections

Offshore wind turbines in shallow waters are predominantly installed using a monopile foundation, onto which a transition piece and wind turbine are attached. Previously, the monopile to transition piece (MP-TP) connection was made using a grouted connection, however, cases of grout failure causing turbine slippage, among other issues, were reported. One solution is to use bolted ring flange connections, which involve using a large number of M72 bolts to provide a firm fixing between the MP-TP. It is in the interest of offshore wind operators to reduce the number of maintenance visits to these wind turbines by maintaining a preload (Fp) level above the minimum requirement for bolted MP-TP connections. The present study focuses on the effect of the tightening sequence on the Fp behaviour of M72 bolted connections. A detailed finite element (FE) model of a seven-bolt, representative segment of a monopile flange was developed with material properties obtained from the available literature. Three analyses were made to examine the effect on Fp after tightening, including the initial Fp level applied to the bolts, the tightening sequence and the effect of an additional tightening pass.

Impact of a Combustor-Turbine Interface on Turbine Vane Aerodynamics and Film Cooling

Optimizing the aero-thermal performance of the combustor-turbine interface is an important factor in enhancing the efficiency of heavy-duty gas turbines. Also, it is a key requirement to fulfill the lifetime in this hottest area of the gas turbine. Typically transition pieces of can combustors induce a highly non-uniform swirling flow at the turbine inlet. In order to better understand the impact of the non-uniform combustor flow at the first stage vanes, a combined experimental and numerical study was carried out. The experimental facility consisted of a high speed linear cascade with four vane passages, including an upstream transition piece, which was representative of a heavy duty gas turbine can combustor-turbine interface geometry. The experiments were conducted at engine representative Mach numbers and film cooling effectiveness measurements were performed at three different blowing ratios. CFD RANS simulations were undertaken using a commercial flow solver. The numerical model was first validated with the experimental data, using inlet traverse 5-hole probe measurements, pressure taps along the airfoil perimeter and oil flow visualization results. The investigation shows that the position of the vane relative to the combustor transition piece has a significant impact on the vane aerodynamics and also film cooling behavior. This understanding was key to a robust first vane aerothermal design of the GT36.