An Advanced Radial-Component Industrial Turbine Engine

Effect of internal heat leakage on the performance of a high pressure ratio centrifugal compressor

The paper deals with the clarification of the unsteady, transonic flow phenomena occurring in the vaned diffuser area of a high pressure ratio centrifugal compressor stage. A description of the test rig, the measurement technique applied and a brief description of the steady 3D code used for the theoretical investigations are given. The measured stage characteristic and the mean diffuser pressure rise as deduced from the 3D calculations are presented. The unsteady flow character in the vaned diffuser region is experimentally resolved with the help of laser measurements performed for the whole diffuser region, from impeller exit up to the vaned diffuser exit. A comparison with the results of a steady state 3D calculation carried out separately for the rotor and diffuser shows good agreement with reality, especially with the flow in the rear part of the diffuser, where the real flow was found to be almost steady.

A three-dimensional Navier-Stokes solver is used to investigate the flow field of a high pressure ratio centrifugal compressor for turbocharger applications. Such a compressor consists of a double-splitter impeller followed by a vaned diffuser. Particular attention is focused on the analysis of the vaned diffuser, designed for high subsonic inlet conditions. The diffuser is characterized by a complex three-dimensional flow field, and influenced by the unsteady interaction with the impeller. Detailed Particle Image Velocimetry (PIV) flow measurements within the diffuser are available for comparison purposes.

Future turbo systems for aircraft engines need very compact geometry, low weight and high efficiency components. The geared turbofan enables the engine designer to decouple the speed of the fan and the LP turbine to combine a low speed fan with a high speed LP turbine. The low pressure turbine is a key component for this concept. The technological challenge is very much driven by the very high low-spool speed. Resulting from also high inlet temperatures the LP turbine needs cooling of the first stage. A new MTU LPT concept for such a high speed turbine has been developed and tested in a turbine rig. The concept consists of a 2 stage turbine for extremely high speed and high stage pressure ratio (ER 2.3). This leads to extra high mechanical loading and an exotic combination of high Mach Numbers (transonic) and very low Reynolds Numbers. In this paper some design features are described. Some elements of the airfoil design were also tested in additional cascade tests. The 2 stage turbine was tested at the Altitude Test Facility of the ILA, Stuttgart. The test setup is described including details of the instrumentation. Test data shows a good turbine performance. Measurements are also compared to 3D CFD, which is used to analyse local effects.

Microturbines have been developed as an important power unit for distributed generation (DG) or distributed energy resource (DER) options [1]. They have been established and are widely used in aircraft and power applications, due to their easy installation, reliability, high performance, multi-fuel capabilities and low emission [2]. However, the aerothermodynamic design of a radial turbine still poses a challenge due to its high rotational speed and high inlet temperature, which influence the centrifugal stress and the rotor structural integrity. This paper presents the numerical investigations on the aerothermodynamic design of the nozzle and the radial inflow rotor for a 600 kW simple cycle gas turbine engine using a One-dimensional Computer FORTRAN Code (OFC) [3], on the grounds of non-dimensional parameters aimed at computational and work time reduction. This program utilizes a one-dimensional solution of flow conditions through the turbine along the meanline. The referred computer program is an effective performance prediction tool mostly in the initial stage of the preliminary design and can be used to quickly investigate and calculate the number of design options prior to any details of the vane and blade geometry. In order to find the most promising design option, a computational fluid dynamics (CFD) simulation has been used to study the performance, the aerothermodynamic design and the flow characteristics of the turbine components. The OFC results were compared with the CFD simulation, a computer program for the design analysis of radial inflow turbines, and analytical results taken from specialized literature showed the results were in agreement.

Computational Fluid Dynamics (CFD) simulations are widely used in mechanical engineering. Although achieving a high level of confidence in numerical modelling is of crucial importance in the field of turbomachinery, verification and validation of CFD simulations are very tricky especially for complex flows encountered in radial turbines. Comprehensive studies of radial machines are available in the literature. Unfortunately, none of them include enough detailed geometric data to be properly reproduced and so cannot be considered for academic research and validation purposes. As a consequence, design improvements of such configurations are difficult. Moreover it seems that well-developed analyses of radial turbines are used in commercial software but are not available in the open literature especially at high pressure ratios. It is the purpose of this paper to provide a fully open set of data to reproduce the exact geometry of the high pressure ratio single stage radial-inflow turbine used in the Sundstrand Power Systems T-100 Multipurpose Small Power Unit. First, preliminary one-dimensional meanline design and analysis are performed using the commercial software RITAL from Concepts-NREC in order to establish a complete reference test case available for turbomachinery code validation. The proposed design of the existing turbine is then carefully and successfully checked against the geometrical and experimental data partially published in the literature. Then, three-dimensional Reynolds-Averaged Navier-Stokes simulations are conducted by means of the Axcent-PushButton CFD® CFD software. The effect of the tip clearance gap is investigated in detail for a wide range of operating conditions. The results confirm that the 3D geometry is correcty reproduced. It also reveals that the turbine is shocked while designed to give a high-subsonic flow and highlight he importance of the diffuser.

Turbine optimization potential to improve automotive Rankine cycle performance

The impeller stall inception in a high-pressure ratio centrifugal compressor with bleed slots was investigated by experiments and unsteady numerical analysis. This study reported that the flow instability of the internal flow was first initiated around the leading edge of the splitter blades and finally resulted in tip leakage flow instability at the leading edge of the main blades. The impeller stall consisted of a stall cell and rotated about 70% of the impeller rotational speed. Furthermore, a pressure disturbance occurs at the splitter blade leading edge before the main blade leading edge. The blockage within the passage of the suction side at the leading edge of the splitter blade was first increased because the leading-edge separation on the suction surface of the splitter blade was enlarged. The blockage within the passage of the pressure side was increased after that of the suction side stopped increasing. Therefore, the blockage development of both pressure and suction passage at the leading edge of the splitter blade was the main cause of the stall inception. Then, a longitudinal vortex was formed near the bleed slots and blocked the mainstream near tip side. Subsequently, the reverse flow region in the inducer was developed and extended to the impeller inlet. Finally, the flow angle to the main blade increased, the spillage occurred at a leading edge of a main blade, and the rotating stall was generated. The cause of impeller instability was the separation of the suction surface of the splitter blade and the formation of a stagnation region on the pressure surface at the splitter leading edge.

This paper aims at carrying out comparative performance analysis of simple and advanced cycles large-scale aero-derivative industrial gas turbines derived from aircraft turbofan engines. The investigation involves technical performances of three large-scale aero-derivative engine cycles based on existing and projected cycles for applications in land based power generation and Combined-Heat-and-Power (CHP). Preliminary design and performance simulation were implemented of a simple cycle (baseline) three-spool 100 MW aero-derivative engine model, intercooled and intercooled/recuperated engine cycles of the same 100 MW nominal power rating. In the analysis, design point and off-design performances of the engine models were established. The results indicate that to a large extent, the advanced engine cycles showed superior performance in terms of thermal efficiency, and fuel flow. In numerical terms, thermal efficiencies of intercooled engine cycle, and intercooled/recuperated engine cycles, over the simple cycle at design point increased by 2.42% and 0.94% respectively, whereas heat rates of these cycles over simple cycle at design point decreased by 2.37% and 0.93% respectively. It is worthy of note that for large-scale aero-derivative gas turbines having power rating of 100 MW and above, intercooled cycle would consume less fuel than intercooled-recuperated and simple cycles. This finding would actually aid good choice of cycle option for large-scale aero-derivative gas turbine designers, manufacturers and users.

In case of small torque air turbine has often been used for the purpose to obtain high speed, for example to drive Gyro, Ultracentrifuge, and Internal-Grinder etc. In this case, usually radial flow impulse type has been adopted for its simple construction. It appears that there are few reports proposed up to this time concerning the turbine of such type with which we can conclude turbine efficiency, and so concerning Parallel nozzles used for it. An investigation was made to present a complete design data applied qualitatively to this type turbine design. The contents of this paper are as follows : (1) The characteristics of the radial flow type air impulse turbine and the best profile of buckets have been presented and checked by experiments. (2) Experiments of the radial flow type air reaction turbine have been given. (3) The author points out that bucket velocity coefficient of the turbines of this type will be about 0·6, and the comparison of the two has been discussed. (4) Results of reaction tests of parallel nozzles have been given, and he presents analytically simple formulae to conclude velocity and discharge coefficient of parallel. nozzles with arbitrary dimensions.

Radial turbines that have a permanent magnet generator rotor in a common shaft have been widely used for small turbines like a microturbine. Though one of the features for those turbo machineries is their capability to be configured with a simplified rotor configuration without gear boxes, thrust generated by the rotor system has to be canceled in the rotor system itself. A simple method is presented for thrust estimation which has been used in development of a 150kW class microturbine. This simple method for thrust estimation was applied to design a similar type of small radial inflow turbine. The turbine rotor was designed to satisfy both high turbine efficiency and low thrust. Experimental results of the prototype turbine load tests showed that the estimated thrust was within the design allowance range at the rated load.

Estimation of the effective nominal power of a photovoltaic generator under non-ideal operating conditions

The current design trend of centrifugal compressors is aimed at obtaining compact machines with high pressure ratios and performance. This leads to high circumferential speeds and lighter components, thus increasing static stresses and dynamic forcing due to rotor/stator interactions. Therefore, the aerome-chanical characterization of these machines is a fundamental step of the design chain. This paper presents an in-depth aerome-chanical study of a high-pressure ratio centrifugal compressor using CFD. Firstly, the computational setup is validated against the experimental data; then, the same numerical setup is used to characterize the aeromechanical interactions between the impeller and the upstream and downstream static components. The URANS simulations are carried out at three different operating points (near stall, design, near choke), and the unsteady flow field is time and space decomposed showing higher forcing functions at near stall and near choke conditions with respect to the design point, as confirmed by the experimental acquisitions. The presented results highlight the importance of extending the aeromechanical evaluations to different operating points in order to avoid unexpected vibration responses when operating at off-design conditions. Finally, by virtue of an acoustic post-processing techniques, the unsteady pressure coupling in the inter-row region is analyzed to discern the main sources of the overall unsteady interaction.

The current drive for high efficiencies and low emissions has resulted in the examination and development of advanced power systems based on complex gas turbine cycles. Reheat and intercooling are two such schemes. The basic objective of introducing intercooling and reheat is to sources of loss in a gas turbine engine are those arising from the turbomachinery and the need to cool the turbine blades. In this paper the concepts of intercooling and reheat for gas turbines are assessed, in a systematic way, using a model that includes the above losses in order to evaluate their effects on the engine performance. Also examined is the choice of the position where intercooling or reheat is implemented which can have a large effect on the engine output. A comparison is made with the simple cycle and it is shown that these schemes show much promise. Some of the development difficulties are also outlined. Intercooling promises large improvements in efficiency over the simple cycle, especially at high pressure ratios. Reheat on the other hand is much more suited to combined cycles.

This paper presents a CFD-based numerical analysis on the potential benefits of non-radial blading turbine for low speed-low pressure applications. Electric turbocompounding is a waste heat recovery technology consisting of a turbine coupled to a generator that transforms the energy left over in the engine exhaust gases, which is typically found at low pressure, into electricity. Turbines designed to operate at low specific speed are ideal for these applications since the peak efficiency occurs at lower pressure ratios than conventional high speed turbines. The baseline design consisted of a vaneless radial fibre turbine, operating at 1.2 pressure ratio and 28,000rpm. Experimental low temperature tests were carried out with the baseline radial blading turbine at nominal, lower and higher pressure ratio operating conditions to validate numerical simulations. The baseline turbine incidence angle effect was studied and positive inlet blade angle impact was assessed in the current paper. Four different turbine rotor designs of 20, 30, 40 and 50° of positive inlet blade angle are presented, with the aim to reduce the losses associated to positive incidence, specially at midspan. The volute domain was included in all CFD calculations to take into account the volute-rotor interactions. The results obtained from numerical simulations of the modified designs were compared with those from the baseline turbine rotor at design and off-design conditions. Total-to-static efficiency improved in all the non-radial blading designs at all operating points considered, by maximum of 1.5% at design conditions and 5% at off-design conditions, particularly at low pressure ratio. As non-radial fibre blading may be susceptible to high centrifugal and thermal stresses, a structural analysis was performed to assess the feasibility of each design. Most of non-radial blading designs showed acceptable levels of stress and deformation.