New compact engine architectures such as pressure gain combustion require ad-hoc turbomachinery to ensure an adequate range of operation with high performance. A critical factor for supersonic turbines is to ensure the starting of the flow passages, which limits the flow turning and airfoil thickness. Radial outflow turbines inherently increase the cross section along the flow path, which holds great potential for high turning of supersonic flow with a low stage number and guarantees a compact design. First the preliminary design space is described. Afterwards a differential evolution multi-objective optimization with 12 geometrical design parameters is deducted. With the design tool AutoBlade 10.1, 768 geometries were generated and hub, shroud, and blade camber line were designed by means of Bezier curves. Outlet radius, passage height, and axial location of the outlet were design variables as well. Structured meshes with around 3.7 million cells per passage were generated. Steady three dimensional Reynolds averaged Navier Stokes (RANS) simulations, enclosed by the k-omega SST turbulence model were solved by the commercial solver CFD++. The geometry was optimized towards low entropy and high power output. To prove the functionality of the new turbine concept and optimization, a full wheel unsteady RANS simulation of the optimized geometry exposed to a nozzled rotating detonation combustor (RDC) has been performed and the advantageous flow patterns of the optimization were also observed during transient operation.

Modern lean burn aero-engine combustors make use of relevant swirl degrees for flame stabilization. Moreover, important temperature distortions are generated, in tangential and radial directions, due to discrete fuel injection and liner cooling flows respectively. At the same time, more efficient devices are employed for liner cooling and a less intense mixing with the mainstream occurs. As a result, aggressive swirl fields, high turbulence intensities, and strong hot streaks are achieved at the turbine inlet. In order to understand combustor-turbine flow field interactions, it is mandatory to collect reliable experimental data at representative flow conditions. While the separated effects of temperature, swirl, and turbulence on the first turbine stage have been widely investigated, reduced experimental data is available when it comes to consider all these factors together.In this perspective, an annular three-sector combustor simulator with fully cooled high pressure vanes has been designed and installed at the THT Lab of University of Florence. The test rig is equipped with three axial swirlers, effusion cooled liners, and six film cooled high pressure vanes passages, for a vortex-to-vane count ratio of 1:2. The relative clocking position between swirlers and vanes has been chosen in order to have the leading edge of the central NGV aligned with the central swirler. In order to generate representative conditions, a heated mainstream passes though the axial swirlers of the combustor simulator, while the effusion cooled liners are fed by air at ambient temperature. The resulting flow field exiting from the combustor simulator and approaching the cooled vane can be considered representative of a modern Lean Burn aero engine combustor with swirl angles above ±50 deg, turbulence intensities up to about 28% and maximum-to-minimum temperature ratio of about 1.25. With the final aim of investigating the hot streaks evolution through the cooled high pressure vane, the mean aerothermal field (temperature, pressure, and velocity fields) has been evaluated by means of a five-hole probe equipped with a thermocouple and traversed upstream and downstream of the NGV cascade.

For solar plants, waste-energy recovery, and turbogenerators, there is a considerable amount of waste energy due to low mass flow rate. Owing to the high specific power output and large pressure ratios across the turbine, a supersonic industrial steam turbine (IST) is able to utilize the waste energy associated with low mass flow rate. Supersonic IST has fewer stages than conventional turbines and a compact and modular design, thus avoiding the excessive size and manufacturing cost of conventional IST. Given their flexible operation and ability to function with loads in the range of 50–120% of the design load, supersonic IST offers significant advantages compared to conventional IST. The strong shock-wave loss caused by supersonic flows can be reduced by decreasing the shock intensity and reducing its influence; consequently, a supersonic IST can reach higher efficiency levels. Considering the demonstrated utility of bowed blades in conventional IST, this paper presents a study of the use of bowed blades in a supersonic IST. For this purpose, first, the shock-wave structure in the supersonic flow field was analyzed and compared with experimental results. Then, four different bowed blades were designed and compared with a straight blade to study the influence of bowed blades on the shock-wave structure and wetness. The results indicate that S-shaped bowing can improve the efficiency of supersonic turbines, and the energy-loss coefficient of the stators can be decreased by 2.4% or more under various operating conditions.

In large modern turbochargers, transonic compressors often constitute the main source of noise, with a frequency spectrum typically dominated by tonal noise at the blade passing frequency (BPF) and its harmonics. Inflow BPF noise is mainly generated by rotor locked shock fronts. Outflow noise, while also dominated by BPF tones, is linked to more complex source mechanisms. Its modal structure and the relationships between sources and modal sound pressure levels (SPL) are less well understood, and its numerical analysis is, in general, significantly more complex than for compressor inflows. To shed some light on the outflow acoustic characteristics of radial machines, transient simulations of a 360 deg model of a radial compressor stage, including its vaned diffuser and volute, were carried out. Four increasingly finer grids were used for this purpose. On all grids, numerical damping had detrimental effects on prediction quality. A simple and mathematically sound method is proposed to account for this damping. With it, the global outflow acoustic power level ( PWL g ) is predicted to within an accuracy of 2 dB of the experimental result on the finest grid. This shows that satisfactory accuracy can be obtained with state-of-the-art computational fluid dynamics (CFD) codes if care is taken with the simulation setup. The simulations are further validated with experimental data from 17 transient wall pressure sensors.

Porous media model computational fluid dynamics (CFD) is a valuable approach allowing an entire heat exchanger system, including the interactions with its associated installation ducts, to be studied at an affordable computational effort. Previous work of this kind has concentrated on developing the heat transfer and pressure loss characteristics of the porous medium model. Experimental validation has mainly been based on the measurements at the far field from the porous media exit. Detailed near field data are rare. In this paper, the fluid dynamics characteristics of a tubular heat exchanger concept developed for aero-engine intercooling by the authors are presented. Based on a rapid prototype manufactured design, the detailed flow field in the intercooler system is recorded by particle image velocimetry (PIV) and pressure measurements. First, the computational capability of the porous media to predict the flow distribution within the tubular heat transfer units was confirmed. Second, the measurements confirm that the flow topology within the associated ducts can be described well by porous media CFD modeling. More importantly, the aerodynamic characteristics of a number of critical intercooler design choices have been confirmed, namely, an attached flow in the high velocity regions of the in-flow, particularly in the critical region close to the intersection and the in-flow guide vane, a well-distributed flow in the two tube stacks, and an attached flow in the cross-over duct.

Flow visualization and heat transfer measurements have been made in a cavity comprising two nonplane disks of 762 mm diameter and a peripheral shroud, all of which could be rotated up to 2000 rpm. “Cobs,” made from a lightweight foam material and shaped to model the geometry of turbine disks, were attached to the center of each disk. Cooling air at flow rates up to 0.1 kg/s entered the cavity through the center of the “upstream” disk and left via holes in the shroud. The flow structure was found to be similar to that observed in earlier tests for the plane-disk case: a source region, Ekman layers, sink layer, and interior core were observed by flow visualization. Providing the source region did not fill the entire cavity, solutions of the turbulent integral boundary-layer equations provided a reasonable approximation to the Nusselt numbers measured on the heated “downstream” disk.

The technology of high-pressure air or hot-gas impingement from stationary shroud supplementary nozzles onto radial outflow compressors and radial inflow turbines to permit rapid gas turbine starting or power boosting is discussed. Data are presented on the equivalent turbine component performance for convergent/divergent shroud impingement nozzles, which reveal the sensitivity of nozzle velocity coefficient with Mach number and turbine efficiency with impingement nozzle admission arc. Compressor and turbine matching is addressed in the transient turbine start mode with the possibility of operating these components in braking or reverse flow regimes when impingement flow rates exceed design.

This paper describes a fully explicit, time marching, corrected viscosity, finite volume method to solve the Euler equations in a cylindrical coordinate system. The rotational character of the incoming flow can be taken into account. On the outflow boundary, a generalized radial equilibrium condition is imposed. Blade rows of complex geometry can be handled. At present, the method has been used to calculate the flow through the nozzle vanes of the VKI low-speed turbine facility. The calculated results show good agreement with the experimental data for the spatial distribution of both the static pressure and the flow angle.

In order to gain an understanding of the conditions inside air-cooled, gas-turbine rotors, flow visualization, laser-doppler anemometry, and heat-transfer measurements have been made in a rotating cavity with either an axial throughflow or a radial outflow of coolant. For the axial throughflow tests, a correlation has been obtained for the mean Nusselt number in terms of the cavity gap ratio, the axial Reynolds number, and rotational Grashof number. For the radial outflow tests, velocity measurements are in good agreement with solutions of the linear (laminar and turbulent) Ekman layer equations, and flow visualization has revealed the destabilizing effect of buoyancy forces on the flow structure. The mean Nusselt numbers have been correlated, for the radial outflow case, over a wide range of gap ratios, coolant flow rates, rotational Reynolds numbers, and Grashof numbers. As well as the three (forced convection) regimes established from previous experiments, a fourth (free convection) regime has been identified.

A previously described, inviscid design technique has been substantially improved to allow the generation of either shock-free, weak-shock or low total pressure loss supersonic rotor cascade designs. Improvements have been introduced in inflow-outflow boundary conditions, imposition of geometric constraints and in shock pressure rise specifications. Calculation examples are presented for precompression type rotor designs.

An integral method to predict the performance of the incompressible, turbulent flow between a rotating disk and a parallel stationary wall, when there is radial outflow of ventilation air, is presented. Using this method, the effects of the inlet conditions of the ventilation air on the core rotation, on the radial velocity profile development, on the radial inflow rates, on the separation streamlines between outflow and inflow regions and on the disk torque coefficient are calculated and presented. The method is general enough to calculate other effects. The important conclusions are: (a) that inlet radial velocity profile should be skewed toward the stationary wall, to reduce radial inflow, and (b) that disk friction can be decreased by increasing the rotation of the ventilation air.

Technological progress requires that lessons learned in initial design efforts be incorporated into the following design effort. One area where improvement can be made is in the design of elastomer seals for containment of breeder reactor cover gases. Liquid metal fast breeder reactor (LMFBR) vessels, unlike PWR’s, operate at a very low pressure. The reactor cover gas may be maintained at a pressure of less than 1 psia. This low pressure level and moderate temperatures of the vessel head-mounted refueling and monitoring equipment permit the use of elastomer seals in this equipment to contain the cover gas. Existing seal designs are based on a conservative design philosophy which was established in the late 1960’s when the FFTF project was in its formative stages. The information gained in the design and development of LMFBR seals during the past decade indicates that some changes in the approach to seal design could result in improved seals and reduced downtime for seal maintenance. This document presents the design approach currently used for cover gas seals, and points out how improvements can be made based on today’s knowledge of breeder reactors and elastomer seals. A suggested new seal design concept is presented which may better meet the LMFBR commercial program needs. The basic assumption used to design cover gas seals conservatively states that elastomer seals, even when new, can leak, and primary consideration must be given to preventing the outflow of radioactive cover gas. The suggested new design concept acknowledges that elastomer seals are effective, and the primary consideration should be one of controlling the gases which permeate through the elastomer seal material. Finally, information which was obtained from the cover gas seal development program for FFTF, which is of value to all LMFBR equipment designers, is summarized.

This paper describes a combined theoretical and experimental study of the heat transfer from an air-cooled gas turbine disk using the model of a plane disk rotating close to a shrouded stator. Numerical solutions of the boundary layer equations are obtained by assuming a modified system geometry, and it is shown that this technique yields adequate estimates of moment coefficients and mean Nusselt numbers for the air-cooled disk. Experimental results show the effect of rotational speeds up to 4000 rev/min, coolant flow rates up to 2 lb/s, stator clearances and shroud clearances up to 2.7 in., on the mean Nusselt numbers for a 30-in-dia disk and its stator.

The effects of rotation and secondary flows on flow in rotating, radial-flow passages are investigated. Pressure and hot-wire measurements are made on the developing flow in a single, rotating radial-outflow channel of simple geometry. The channel has straight radial walls, and is of constant height. It is double shrouded, has a total angle of 15 deg, and an area ratio of 3.1:1. The results indicate the formation of a large wake region on the suction (trailing) side of the passage, at high flow rates. At low flow rates, a pressure-side eddy was observed.

This paper describes an experimental study of an air-cooled gas turbine disk using the model of a disk rotating near a shrouded stator. Measurements of pressure distribution, frictional moment, and the cooling air flow necessary to prevent the ingress of hot gases over the turbine disk are described for a range of rotational speeds, mass flow rates, and different geometries. The pressure distribution is shown to be calculable by the superposition of the pressure drop due to the shroud and the unshrouded distribution. Moment coefficients are shown to increase with increasing mass flow rate and decreasing shroud clearance, but are little affected by the rotor/stator gap. Applying Reynolds analogy to the moment coefficients, it is estimated that heat transfer from the rotor will be controlled primarily by rate of radial cooling flow at low rotational Reynolds numbers, and will be governed primarily by Reynolds number at large rotational speeds.