In order to cover non-uniform blade tip gaps within a compressor mean line performance program a ... more In order to cover non-uniform blade tip gaps within a compressor mean line performance program a method is developed that derives a generic correlation that converts a discretized gap of any shape into an equivalent constant clearance applicable for common loss models. This method is demonstrated with the example of the NASA Rotor 37. For this purpose, the efficiency losses of uniform gap geometries are compared with those of non-uniform gaps of the same rotor. The efficiency information is provided by flow simulation studies. Geometric weighting factors have been determined for a correlation function, so that it is now possible to infer uniform from non-uniform gaps. This enables a mean line representation of blades with non-uniform tips. In the example of the NASA Rotor 37, the standard deviation of the correlation error of the considered member is below the measurement uncertainty for experimental determination of the gap size. Validation cases confirm this result.
Zenodo (CERN European Organization for Nuclear Research), May 7, 2018
A numerical aeroelastic assessment of a highly loaded high pressure compressor exposed to flow di... more A numerical aeroelastic assessment of a highly loaded high pressure compressor exposed to flow disturbances is presented on this paper. The disturbances originate from novel, inherently unsteady, pressure gain combustion processes, such as pulse detonation, shockless explosion, wave rotor or piston topping composite cycles. All these arrangements promise to reduce substantially the specific fuel consumption of present-day aeronautical engines and stationary gas turbines. However, their unsteady behavior must be further investigated to ensure the thermodynamic efficiency gain is not hindered by stage performance losses. Furthermore, blade excessive vibration (leading to high cycle fatigue) must be avoided, especially under the additional excitations frequencies from waves traveling upstream of the combustor. Two main numerical analyses are presented, contrasting undisturbed with disturbed operation of a typical industrial core compressor. The first part of the paper evaluates performance parameters for a representative blisk stage with high-accuracy 3D unsteady Reynolds-averaged Navier-Stokes computations. Isentropic efficiency as well as pressure and temperature unsteady damping are determined for a broad range of disturbances. The nonlinear harmonic balance method is used to determine the aerodynamic damping. The second part provides the aeroelastic harmonic forced response of the rotor blades, with aerodynamic damping and forcing obtained from the unsteady calculations on the first part. The influence of blade mode shapes, nodal diameters and forcing frequency matching is also examined.
In the present paper, an experimental investigation on the application of high-speed sweeping jet... more In the present paper, an experimental investigation on the application of high-speed sweeping jet actuators was carried out. It examines the feasibility of a super sonic exit jet to control boundary layer separation caused by an adverse pressure gradient in a compressible Mach number regime. The first part of the investigation includes a thorough assessment of the used sweeping jet actuator by three dimensional numerical simulations using OpenFoam ®. A validation of the results was done by means of schlieren visualization within the actuator and dynamic pressure measurements at the actuator outlet. The outlet signal of the actuator features a characteristic switching frequency of f s ≈ 1200 Hz and a peak Mach number of M a ≈ 1.4 which was evident in the experiments and the CFD simulation. The supply pressure ratio of the actuator was set to P R = 3.6 using compressed air. The second part of the paper presents the results of the wind tunnel experiments using a test section with the geometry of a half-diffusor ramp equipped with two sweeping jet actuators. Inflow conditions of Mach numbers ranging from M a ∞ = 0.3 to M a ∞ = 0.8 were investigated. The conducted wind tunnel experiments include detailed oil flow surface visualizations and mean surface pressure measurements in order to demonstrate the impact of the active flow control system. By varying the mass flow rate of the actuator system different operating points were studied. The presented results indicate the positive effect of active flow control regarding the pressure rise over the ramp in a compressible Mach number regime. The separated flow region was significantly reduced.
Much of academic research on gas turbines is conducted at component level. Promising concepts suc... more Much of academic research on gas turbines is conducted at component level. Promising concepts such as enhanced blade design or active flow control are often investigated on isolated domains, e.g. single compressor stages. Possible interactions with adjacent components, such as effects on secondary power and air provision or the thermodynamic cycle are rarely or just subsidiary considered. Nevertheless, the early inclusion of those effects is crucial for a reliable assessment of novel concepts' benefits. For example, the reduction of aero engine component weight or the increase of turbo-component stage efficiency is intangible regarding the quantified benefit on system level. This gap can be closed by a technique called component zooming. Here, a gas turbine performance model representing the cycle and components basic interaction is coupled with more detailed models which are capable of specific investigations on component level. This paper presents a suitable architecture of frameworks for component zooming in general. In order to maintain high applicability especially for academic purposes, focus is put on flexibility in manifold sense: Adaptability to arbitrary concepts and zoomed components, expandability of available workflows, expandability to new disciplines, exchangeability of used software (e.g. commercial or in-house) and portability to common platforms. Various examples will demonstrate the capabilities of component zooming performed with such a framework. The first case discusses zooming on the secondary air system (SAS) of gas turbines. Another case demonstrates the coupling of engine performance with models for an initial assessment of aero engine components weight reduction benefit, e. g. realized by tandem blades or flow actuation in duct components. The resulting change in overall mass of the engine leads to basic, pragmatic options for aircraft operators.
Future engines tend to have increasingly higher bypass ratios in order to reduce fuel consumption... more Future engines tend to have increasingly higher bypass ratios in order to reduce fuel consumption. This goes along with a decrease of the fan pressure ratio which may result in an unchoked nozzle. Thus the fan working line drops whereby the efficiency of the fan can be reduced. To avoid this effect variable area thrust nozzles can be utilized. For the evaluation of the potentials of variable thrust nozzles with regard to effects on the overall system, a flight mission model of a representative narrow body aircraft is developed based on published data. On the basis of the equations of motion of each individual flight phase, the mission fuel consumption is calculated. A validation of the nominal case on the basis of EUROCONTROL data confirms the accuracy of the results. The nominal case is opposed by an engine with a variable thrust nozzle, which fixes the working line of the fan to the peak efficiency line. It is shown that the fuel savings per flight hour are most relevant on short-haul flights. This is explained by the relatively high proportion of operating points in the part load range, i.e. those points which deviate significantly from the respective peak efficiency point. Furthermore, it is shown that the potential savings of variable thrust nozzles are reduced with increasing bypass ratio, since the reference efficiency increases as well. So the application of such a nozzle modification in future engines can indeed help to stabilize the fan's working line, but not necessarily to operate the aircraft more cost-efficiently.
Volume 5: Manufacturing Materials and Metallurgy; Ceramics; Structures and Dynamics; Controls, Diagnostics and Instrumentation; Education, Jun 2, 1998
The subject of this paper is the development and first analysis of the performance of a new boost... more The subject of this paper is the development and first analysis of the performance of a new booster bleed valve control logic to improve the surge margin of the booster compressor in the new BR715 turbofan. The booster compressor is driven by the LP spool of the two spool engine, comprising an aerodynamically tuned system out of fan root and booster compressor. The flow delivered at the exit of the booster compressor must be accepted by the following HP compressor. This matching is not always possible, since the rates of change of the spool speeds during transients behave differently as do pressure rise and mass flow. Hence the transient handling of the booster compressor is difficult with respect to its surge stability. To aid in this situation it is a common practice to position a booster bleed valve after the booster compressor, which passes bleed air into the bypass duct. The BR715 jet engine takes advantage of a newly developed logic driving this valve. Monitoring input parameters such as spool and flight speeds as well as altitude, the logic provides the capability of appropriately positioning the valve, thereby maintaining an optimum performance of the engine throughout the whole operating range. Furthermore the logic allows for reaction on thrust reverser deployment or more serious events such as surge and foreign object damage. Much effort has been put into the development of the newly designed transient valve logic, which in its main part is based on the rate of change of the pressure at the exit of the HP compressor. This logic ensures a sufficient surge margin in case of deceleration or reslams of the engine. A simulation of the logic was set up, connected to a synthesis program for the two spool engine and used for the development. After this basic testing, the logic was implemented in the BR715 engine, which first ran in April 1997. The actual data from basic handling tests performed with this engine were analysed and checked against simulation results. The paper describes the logic in detail and contains first results. They prove a proper working of the logic and show the strong effect of the booster bleed valve.
Tradeoff Assessments for Part Load Controlled Cooling Air in Stationary Gas Turbines
Journal of engineering for gas turbines and power, Sep 25, 2020
The demand for flexible part load operation of stationary gas turbines requires the simultaneous ... more The demand for flexible part load operation of stationary gas turbines requires the simultaneous design for sufficient efficiency and lifetime. Both can be addressed by the secondary air system. This paper presents investigations on the concepts of cooling air reduction in off-design, aiming for tradeoffs between fuel burn and turbine blade life. The considered lifetime mechanisms are creep and oxidation. In addition, the effects on emissions from the combustion are outlined. The reference gas turbine is a generic gas turbine in the 300 MW power output segment. The focus is on the first two stages of the four-stage turbine. All simulations are performed by application of a coupled model that essentially connects gas turbine performance with a secondary air system network model. This coupled model is now extended with blade life evaluation and emission models. The results contain tradeoffs for operating points at base and part load. For example, the combined cooling air control of stage 1 rotor blade and stage 2 vane offers savings up to 0.5% fuel flow at 60% of base load in a combined cycle application. This saving is at the expense of creep life. However, some operating points could even operate at higher blade temperatures in order to improve life regarding hot corrosion. Furthermore, generic sensitivities of controlled secondary air supply to cooling layers and hot gas ingestion are discussed. Overall, the presented trades mark promising potentials of modulated secondary air system concepts from a technical point of view.
The design of new stationary gas turbines and development of upgrades for existing respectively i... more The design of new stationary gas turbines and development of upgrades for existing respectively is facing challenges regarding part load operation. The demands for high overall efficiency and compliance with legal requirements depend on the design of cooling air circuits among others. The design of an optimized secondary air distribution at both base load and part load as well as the consideration of different ambient conditions requires conceptual studies and hence appropriate models. This paper introduces the holistic model of a literature based generic stationary gas turbine, which essentially couples a gas turbine performance synthesis model with a more detailed secondary air system (SAS) network model. Extended with additional models such as evaluation of blade and vane material temperatures T mat , it allows for comparative off design studies with uncontrolled and controlled turbine cooling air circuits. The presented studies here first focus on margins of T mat with base load condition as benchmark. The subsequent exploitation of these margins is limited by the fundamental requirements of hot gas ingestion at common rim seal configurations. Either way, the reduction of cooling air at part load is beneficial in terms of fuel flow reduction: vane cooling air control results in up to 0.12% of fuel flow reduction at part load operation.
Prediction of Unsteady 2D-Flow in Turbomachinery Bladings
Springer eBooks, 1993
A numerical algorithm for calculating the unsteady flow in turbomachinery bladings is presented. ... more A numerical algorithm for calculating the unsteady flow in turbomachinery bladings is presented. The unsteadiness can be caused by oscillating cascades and incoming wakes of preceding blade rows simultaneously. The analysis of the time-dependent flow field yields the resulting blade forces and moments on the described profile.
The aim of the present paper is to improve the physical understanding of discrete prestall flow d... more The aim of the present paper is to improve the physical understanding of discrete prestall flow disturbances developing in the tip area of the compressor rotor. For this purpose, a complementary instrumentation was used in a single-stage axial compressor. A set of pressure transducers evenly distributed along the circumference surface mounted in the casing near the rotor tip leading edges measures the time-resolved wall pressures simultaneously to an array of transducers recording the chordwise static pressures. The latter allows for plotting quasi-instantaneous casing pressure contours. Any occurring flow disturbances can be properly classified using validated frequency analysis methods applied to the data from the circumferential sensors. While leaving the flow coefficient constant, a continuously changing number of prestall flow disturbances appears to be causing a unique spectral signature, which is known from investigations on rotating instability. Any arising number of disturbances is matching a specific mode order found within this signature. While the flow coefficient is reduced, the propagation speed of prestall disturbances increases linearly, and meanwhile, the speed seems to be independent from the clearance size. Casing contour plots phase-locked to the rotor additionally provide a strong hint on prestall disturbances clearly not to be caused by a leading edge separation. Data taken beyond the stalling limit demonstrate a complex superposition of stall cells and flow disturbances, which the title “prestall disturbance” therefore does not fit to precisely any more. Different convection speeds allow the phenomena to be clearly distinguished from each other. Furthermore, statistical analysis of the pressure fluctuations caused by the prestall disturbances offer the potential to use them as a stall precursor or to quantify the deterioration of the clearance height between the rotor blade tips and the casing wall during the lifetime of an engine.
On the Effect of Volute Tongue Design on Radial Turbine Performance
With an increasing need for gas turbines with rather low flow rates in many industrial applicatio... more With an increasing need for gas turbines with rather low flow rates in many industrial applications, e.g. decentralized power generation, aircrafts or automotive turbochargers, the development of small size radial turbines becomes more and more important. A major step in the development of a radial turbine stage is the preliminary design, which is the definition of basic geometrical features and the calculation of general turbine flow parameters at the design point and within the operating range. These are mainly the rotational speed, the expansion ratio, the flow rate and in particular the expected turbine efficiency. In a radial turbine stage, the volute component delivers the flow to the rotor wheel and according to the geometrical form it defines major flow parameters like the mass flow parameter or the absolute rotor inlet flow angle. Amongst others, the way the flow enters the turbine wheel represents one of the most important loss generating factors. Thus, on the one hand an approach is necessary for the calculation of the optimum rotor inlet flow angle, in order to avoid dispensable losses due to secondary flow in the turbine wheel region. On the other hand, the volute tongue generates flow non-uniformity which has an effect on the overall circumferential averaged rotor inlet flow angle. Furthermore, the local flow pattern downstream of the volute tongue can generate suboptimal flow conditions for the turbine wheel. Hussain and Bhinder [1] measured the flow field at the outlet of a vaneless volute at different circumferential positions and detected a variation of the outlet angle of about Δα = 10°. The authors conclusion was, that the influence on the stage performance of flow non-uniformity generated by the volute could exceed the one of pressure losses through the volute. In this paper, the effect of different geometrical volute parameters on the flow condition especially at the turbine wheel inlet area is investigated. Experimental data of the influence of different volute tongue geometries on the flow field is difficult to generate. Hence, comprehensive numerical investigations are made using steady 3D-CFD calculations of the turbine volute as well as calculations of complete turbine stages including a turbine wheel geometry. Based on the numerical results, a design guideline is developed to estimate the influence of the geometric volute parameters on the flow and to raise the quality of the preliminary design process.
Axial compressors in aero engines are prone to suffering a breakdown of orderly flow when operati... more Axial compressors in aero engines are prone to suffering a breakdown of orderly flow when operating at the peak of the pressure rise characteristic. The damaging potential of separated flows is why a safe distance has to be left between every possible operating point and an operating point at which stall occurs. During earlier investigations of stall inception mechanisms, a new type of prestall instability has been found. In this study, it could be demonstrated that the prestall instability characterised by discrete flow disturbances can be clearly assigned to the subject of "Rotating Instabilities". Propagating disturbances are responsible for the rise in blade passing irregularity. If the mass flow is reduced successively, the level of irregularity increases until the prestall condition devolves into rotating stall. The primary objective of the current work is to highlight the basic physics behind these prestall disturbances by complementary experimental and numerical investigations. Before reaching the peak of the pressure rise characteristic flow, disturbances appear as small vortex tubes with one end attached to the casing and the other attached to the suction surface of the rotor blade. These vortex structures arise when the entire tip region is affected by blockage and at the same time the critical rotor incidence is not exceeded in this flow regime. Furthermore, a new stall indicator was developed by applying statistical methods to the unsteady pressure signal measured over the rotor blade tips, thus granting a better control of the safety margin.
Validation and Development of Loss Models for Small Size Radial Turbines
Today an increasing need for gas turbines with extremely low flow rates can be noticed in many in... more Today an increasing need for gas turbines with extremely low flow rates can be noticed in many industrial sectors, e.g. power generation, aircraft or automotive turbo chargers. For any application it is essential for the turbine to operate at best possible efficiency. It is known that for turbines the specific optimum achievable power output decreases with smaller size. A major contribution for this reduction in efficiency comes from the relative increase of aerodynamic losses in smaller turbine stages. In the early turbine design stage, easy and fast to use two-dimensional calculation codes are widely used. In order to produce qualitatively good results, all of these codes contain a diversity of loss models that more or less exactly describe physical effects which generate losses. It emerges to be a real problem that most of these empirical models were derived for rather large scale turbo machines and that they are not necessarily suitable for application to small turbines. In this paper many of the commonly known and well established loss models used for the preliminary design of radial turbines were collected, reviewed, and validated with respect to their applicability to small-size turbines, i.e. turbines of inlet diameter smaller than 40 mm. Comprehensive numerical investigations were performed and the results were used to check and verify the outcome of loss models. Based on the results, loss models have been improved. Furthermore, new correlations were developed in order to raise the quality of loss prediction especially for the design of small-size turbines. After receiving an optimum set of loss prediction models, all of them were implemented into a two-dimensional solver program for the analytical iterative solution of a complete turbine stage. Hence a powerful tool for preliminary radial turbine design has been created. This program enables the user to analytically evaluate the effects of changing key design properties on performance. These are amongst others the optimum rotor inlet flow angle according to the slip-factor definition, the value of flow deviation, and hence the optimum blade outlet angle for a minimum adverse flow-swirl at turbine outlet. Complementarily the turbine key performance indicators, e.g. pressure ratio, power output, rotational turbine speed, and mass flow can be calculated for optimum efficiency of a given turbine geometry. The paper presents the most important loss models implemented in the new code and weights their relative importance to the performance of small size radial turbines. The data acquisition was done using the new code itself as well as accompanying full 3D CFD calculations.
The basic concepts and advantages of more/all electric aircraft (M/AEA) are briefly addressed. Th... more The basic concepts and advantages of more/all electric aircraft (M/AEA) are briefly addressed. The combined starter/generator (CS/G) system is introduced as a key technology to enable M/AEA. Some important performance requirements for CS/G system are obtained. Based on these requirements, a high speed switched reluctance machine (SRM) is designed to operate as a starter/generator. The entire design process is mainly divided into two stages: electromagnetic design and thermal design. In electromagnetic design stage, the electromagnetic structure and dimensions of the machine and the number of phase winding turns per pole are obtained; the topology and main technical details of the converter are briefly introduced as well. In thermal design stage, a liquid-cooling system is designed based on the thermal analysis of the machine. In the end, the performances of the designed SRM are basically verified by simulation. To get high performances, the exciting angles are optimized in two different operating modes respectively, and the optimized performances in the motoring mode are given as well.
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Papers by Dieter Peitsch