Single crystal blades used in high pressure turbine bladed disks of modern gas-turbine engines exhibit material anisotropy. In this paper the sensitivity analysis is performed to quantify the effects of blade material anisotropy orientation on deformation of a mistuned bladed disk under static centrifugal load. For a realistic, high fidelity model of a bladed disk both: (i) linear, and (ii) non-linear friction contact conditions at blade roots and shrouds are considered. The following two kinds of analysis are performed: (i) local sensitivity analysis, based on first order derivatives of system response w.r.t design parameters, and (ii) statistical analysis using polynomial chaos expansion. The polynomial chaos expansion is used to transfer the uncertainty in random input parameters to uncertainty in static deformation of the bladed disk. An effective strategy, using gradient information, is proposed to address the “curse of dimensionality” problem associated with statistical analysis of realistic bladed disk.

Pressure ratios of modern high pressure radial compressors tend to increase along with pressure fluctuations and the excitation potential on the impellers. The vibrational interactions between side cavities, filled with high pressure fluid, and the impeller structure play an important role in designing a machine for reliable operation. However, they are not yet fully understood. Vibrations at frequencies that have been uncritical at lower pressure levels could become critical at a higher pressure level. Additionally, coupling effects between fluid and structure are becoming stronger at higher fluid densities. For a safe and reliable design, the excitation and the damping mechanism of coupled modes has to be better understood. To understand the interaction, especially regarding the damping behavior, of coupled structure and acoustic modes, a comprehension of the behavior of the uncoupled or weakly coupled modes is required. The structural damping ratio is very small and it has been analyzed in existing literature extensively. The damping behavior of uncoupled acoustic modes, however, is not yet well investigated. This paper focuses on the damping behavior of acoustic modes that are weakly coupled to structure modes. Measurement results gathered at the aeroacoustic test rig at the University of Duisburg-Essen are presented. The results show the influence of fluid pressure variations on the damping behavior of acoustic modes. Therefore, the response functions of some selected acoustic modes are evaluated with the Peak-to-Peak method. In general, the damping decreases with increasing fluid pressure. Furthermore, a relationship of the damping ratio, the kinematic viscosity, and the natural frequency of the acoustic modes has been detected.

In the context of an engine design calculation, isentropic or polytropic efficiencies of turbomachinery components are assumed at the outset of the cycle analysis and their values are updated or validated following the aerodynamic design of the components. In the present paper, aerodynamic design calculations of axial-flow compressors and turbines are directly integrated into the corresponding performance component models. This creates a consistent, single-step preliminary design and performance modelling process using a relatively small number of physical and geometric inputs. The aerodynamic design for establishing a component’s overall efficiency is accomplished through a mean-line, stage-by-stage approach where the stagewise isentropic efficiency is calculated employing either loss or semi-empirical correlations. From this process, the stagewise flow annulus radii are also obtained and are used to axially size the component stages assuming the blade aspect ratio and axial gapping distributions. The component flowpath geometry is then produced by simply “stacking” axially the component stages. The developed method is validated against publicly available data for a high-pressure compressor and a low-pressure turbine. Finally, the effectiveness of the method is demonstrated by considering the multi-point design of a High Bypass Ratio Geared Turbofan Engine with bypass Variable Area Nozzle.

The effects of high shaft power offtake in a direct drive, a geared drive, and a novel turbofan configuration are investigated. A design and off-design performance analysis shows the configuration specific limitations and advantages. The more electric aircraft (MEA) concept promises to offer advantages with respect to aircraft performance, maintenance and operating costs. The engines for the MEA concept are based on conventional turbofan architectures. These engines are designed for significantly increased shaft power offtake that is required by the airframe, and the shaft power is usually taken off the high-pressure spool. This can impair the off-design performance of the engine and lead to compromises during engine design and to operability limitations. Taking the power off the low-pressure spool mitigates some of the problems but has other limitations. In this work, an alternative novel turbofan architecture is investigated for its potential to avoid the problems related to high shaft power offtakes. This architecture is called the dual drive booster because it uses a summation gearbox to drive the booster from both the low- and high-pressure spool. The shaft power, if taken off the booster spool, is effectively provided by both the low- and high-pressure spools, which allows the provision of very high power levels. This new concept is benchmarked against a two-spool direct drive and a geared drive turbofan. Furthermore, it is described, how the new architecture can incorporate an embedded motor generator. The presented concept mitigates some of the problems which are encountered during high power offtake in conventional configurations. In particular, the core compressors are less affected by a change in shaft power offtake. This allows higher power offtakes and gives more flexibility during engine design and operation. Additionally, the potential to use the new configuration as a gas turbine-electric hybrid engine is assessed, where electrical power boost is applied during critical flight phases. The ability to convert additional shaft power is compared with conventional configurations. Here, the new configuration also shows superior behavior because the core compressors are significantly less affected by power input than in conventional configurations. The spool speed and its variation is more suitable for electrical machines than in conventional configuration with low-pressure spool power transfer. The dual drive booster concept is particularly suited for applications with high shaft power offtakes and inputs, and should be considered for propulsion of more electric aircrafts.

Flows in an intermediate turbine duct (ITD) connecting high-pressure turbines (HPT) and low-pressure turbines (LPT) are highly complex, influenced by the upstream HP turbine flow structures. Non-uniformities originating from the duct with struts of different sizes also affect the LPT inflow conditions, resulting in reduced efficiency. The goal of this paper is to provide detailed understanding of the flow physics and loss mechanisms within the ITDs for highly efficient ITD designs. Steady and unsteady numerical simulations of flows through the ITDs in the presence of HP blade and LP vane were conducted. Effects of upstream HP blade on flow fields and loss characteristics within the ITDs are explored. The generation and propagation of wake and secondary flows through the whole configuration is described, including the fast Fourier transformation (FFT) analyses of the flow in the ITD. Results from the numerical simulations show complex flow patterns resulted from blade-strut-vane flow interactions in a high-endwall-angle duct, which are not obtainable from ITD-only simulations. Moreover, the ITD has a strong amplifying effect on the distorted inflow, and the inflow with the upstream wake and secondary flows introduces a high loss area along the casing at ITD exit. Detailed results are presented and discussed for the flow physics and loss mechanisms within the ITD.

The blockerless cascade thrust reverser is one of the innovative thrust reverser systems, which replaces the traditionally mechanical blocker door with the aerodynamic blocker door by high-pressure secondary injection, thus significantly reduces the nacelle weight and the complexity of the actuator, and especially suitable for high-bypass-ratio turbofan engine. In order to obtain the optimum performance of a blockerless cascade thrust reverser system and provide the guidance for the design of the blockerless cascade thrust reverser system, a blockerless cascade thrust reverser system was studied in this paper based on the Response Surface Method (RSM), focusing on the effect of different geometric and aerodynamic parameters on the thrust reverser performance. Results show that the secondary injection with high pressure forms the blockage effect to the fan flow, then forces the fan flow to deflect and discharge from the cascade window, realizing the reverse thrust. The thrust reverser performance is mainly affected by fan pressure ratio (FPR), secondary flow pressure ratio (SPR), secondary injection position (X jet ), secondary injection angle (α jet ) and cascade installation angle (β), and the dominated factors are FPR, SPR and X jet . According to the obtained response equation of the thrust reverser performance, the relationship between reverse thrust efficiency and various parameters are clearly described, and performance of thrust reverser can be quickly evaluated. Significant interaction effects exist between different two factors, which must be taken into consideration in the design process of the blockerless cascade thrust reverser system, especially for the interaction effect between FPR and X jet , interaction effect between FPR and β. Optimization design with objective of maximum reverse thrust was carried out to determine the best parameter settings, and reverse thrust ratio η Trev of 60% is achieved under the constraint of the secondary flow ratio.

Maintenance costs are a substantial contributor to airline operating costs. In this context, understanding, analyzing, and predicting engine performance deterioration is crucial. While diagnostic methods for analyzing the current module and overall engine condition are established in state-of-the-art engine condition monitoring (ECM) systems, deterioration modeling and prognosis are still fields of research. The identification of the contribution of deterioration mechanisms, such as fouling, erosion, and abrasion, to the in-service deterioration poses a key challenge in this area. This paper focuses on a top-down approach for the high pressure compressor (HPC) module. The selected approach is to quantify the contribution of individual deterioration mechanisms to the overall HPC efficiency deterioration based on in-flight measurements. This is accomplished by first using the in-flight measurements to analyze the HPC efficiency loss. Then, the resulting time series of the analyzed HPC efficiency loss are preprocessed. Finally, models of the deterioration mechanisms are fitted to the preprocessed time series. The deterioration models are chosen based on literature references to the respective deterioration mechanisms. As multiple influencing factors affect the deterioration mechanisms, a fleet analysis is conducted to select the model inputs. The fitting process involves a parametric nonlinear regression problem. The outcome is an estimation of the evolution of the deterioration mechanisms over time. This methodology is used to evaluate all available in-service engines of the same type and fleet and to define a fleet model. In the final step, benefits and limitations of the fleet model are investigated.

Centrifugal blowers are widely used in automotive, heating, ventilating, air-conditioning and other industrial purposes. In fact, they allow high-pressure increase to a moderately high mass flow rate, despite the reduced overall dimensions of the system. Numerous authors have worked to increase the fluid dynamic efficiency of the impeller and the volute. However, the recent sound emission standards have imposed tighter constraints, making the noise reduction one of the most challenging target for the design. The reduction of the tonal noise is the main concern for these blowers because the noise at the first harmonic is clearly distinguishable in the noise spectrum. Techniques to reduce this peak generally rely on experimental measurements in aeroacoustics laboratory. In this work, a CFD procedure has been developed to accurately predict the tonal noise in radial bowers. The approach has been validated using real geometries and experimental data. Various turbulence models have been tested to find the best results without the use of excessive computational resources. Moreover, unsteady simulations of the 3D blower have been carried out to analyze the influence of the main geometric parameters on the tonal noise reduction. A parametric design code developed by the authors have been used to change the geometry in order to identify the effect of the main geometrical design parameters on both fluid dynamic and aeroacoustics performance.

Intercooled recuperated turbofan engines with high bypass ratio are becoming a research focus in recent years due to its advantages of relatively better fuel economy, lower emission and noise characteristic. The re-heater can recover waste heat in the exhaust gas downstream of the low pressure turbine to reduce the specific fuel consumption, and the intercooler can improve compression ability of the compressors with sufficient temperature difference between the high pressure compressor and the low pressure turbine. An optimal pressure ratio split is often sought to maximize the effect of the intercooler on improving the compression ability of the compressors. To determine an optimal pressure ratio split, different combinations of pressure ratio between high and low pressure spools must be calculated, and this requires huge amount of work with the traditional method to achieve the suitable cycle selections. In this paper, theoretic thermodynamic analysis is carried out to derive an explicit solution of the optimum pressure ratio split for maximizing the efficiency of the whole compression path. The effects of different variables on the optimum pressure ratio split are investigated according to the correlated variables in the solution function. A comparison calculation is also made to validate the effectiveness and accuracy of the explicit solution. The results show that the optimum pressure ratio split can be achieved with the derived solution function, which will significantly simplify the process of the cycle parameter selection.

The density of Ceramic Matrix Composite (CMC) materials is approximately 1/3 the density of metals currently used for High Pressure Turbine (HPT) blades. A lower density, and consequently lower centrifugal stresses, increases the feasibility of shrouding HPT blades. Shrouding HPT blades improves aerodynamic efficiency, especially for low aspect ratio turbine blades. This paper explores aerodynamic and structural issues associated with shrouding HPT rotor blades. Detailed Navier-Stokes analysis of a rotor blade showed that shrouding improved blade row aerodynamic efficiency by 1.3%, when the clearance was 2% of the blade span. Recessed casings were used. Without a shroud the depth of the recess equaled the clearance. With a shroud the recess depth increased by the shroud thickness, which included a knife seal. There was good agreement between the predicted stage efficiency for the unshrouded blades and the experimentally measured efficiency. Structural analysis showed a strong interaction between stresses in the shroud and peak stresses at the hub of the blade. A thin shroud of uniform thickness only moderately increased maximum blade stress, but there were very high stresses in the shroud itself. Increasing shroud thickness reduced stresses in the shroud, but increased blade stresses near the hub. A single knife seal added to the thin shroud noticeably decreased maximum shroud stress, without increasing maximum blade stress. Maximum stresses due to pressure loads and combined pressure and centrifugal loads were nearly the same as the maximum stresses for individual pressure or centrifugal loads. Stresses due to a 100K temperature difference across the blade walls were much lower than centrifugal or pressure load stresses.

The article describes a refining method for a fuel pump of rocket powerful turbo-pump unit by the joint usage of mathematical optimization software IOSO, meshing complex NUMECA and CFD complex ANSYS CFX. The optimization software was used for automatic change of the geometry of low-pressure impeller, transition duct and high-pressure impeller to find the optimal design. It was mandatory to keep the original variant of the remaining parts of the pump. For this reason, only geometrical parameters of the blades were varied without changing the contours of the pump meridional flow part. The investigated pump consists of five parts: inlet duct, low-pressure screw centrifugal stage, transition duct, high-pressure screw centrifugal stage and volute outlet duct. The pump main parameters with water as the working fluid (based on experiment data) were the following: high-pressure stage rotor speed was 13300 rpm; low-pressure rotor speed was 3617 rpm by gearbox; inlet total pressure was 0.4 MPa; outlet mass flow was 132.6 kg/s at the nominal mode. Creation of vane unit mesh (rotors and stator transition duct) was performed using NUMECA AutoGrid5. Sector models were used for the calculation simplification. The flow around only one blade or screw was considered. Setting up and solution of the task were carried out in the ANSYS CFX solver. Comparison of calculated characteristics of the basic pump with the experimental data was performed before the optimization. The analysis of characteristics for the obtained optimized pump geometry was carried out. It was found that pump with optimized geometry has greater efficiency in comparison with the original pump variant. The obtained reserve can be used to boost the rocket engine, and/or to reduce the loading of the main turbine, which operates in aggressive oxidizing environment.

Endwalls impose a challenge to cool because of the complex system of secondary flows and separation lines disrupting surface film coolant coverage. The interaction of film cooling flows with secondary flow structures is coupled. The momentum exchange of the film coolant with the mainstream affect the formation the secondary flows, which in turn affect the coolant coverage. Therefore, to develop better endwall cooling schemes, a good understanding of passage aerodynamics as affected by interactions with coolant flows is required. This study presents experimental and computational results for cascade representing the first stage nozzle guide vane of a high-pressure gas turbine. The cascade is subsonic, linear, and stationary with an axisymmetrically-contoured endwall. Two cooling flows are simulated; upstream combustor liner coolant-in the form of an aero-thermal profile simulated in the approach flow and endwall slot film cooling, which is injected immediately upstream of the passage inlet. The experiment is run with engine representative combustor exit flow turbulence intensity and integral length scales, with high turbine passage exit Reynolds number of 1.61 × 10 6 . Measurements are performed with various slot film cooling mass flow rate to mainstream flow rate ratios (MFR). Aerodynamic effects are documented with five-hole probe measurements at the exit plane. Varying the slot film cooling MFR results in minimal effects on total pressure loss for the range tested. Vorticity distributions show a very thin, yet intense, cross-pitch flow on the contoured endwall side. Coolant distribution fields that were previously presented for the same cascade are discussed in context of the aerodynamic measurements. A coolant vorticity parameter presenting the advective mixing of the coolant due to secondary flow vorticity is introduced. This parameter gives developers a new prospective on aerodynamic-thermal performance associated with cooled turbine endwall. The numerical study is conducted for the same test section geometry and is run under the same conditions. The applicability of using RANS turbulence closure models for simulating this type of flow is discussed. The effects of including the combustor coolant in the approach flow is also briefly discussed in context of the numerical results.

This study investigates the behavior of a turbopump assembly during critical cavitation of the propellant pumps in the upper rocket engine of the Korea Space Launch Vehicle-II. Turbopumps operate under conditions involving low pressure at the pump inlet and high rotational speeds to allow for a lightweight design. This severe environment can easily cause cavitation to occur in the pump. This cavitation can then cause the pump operation to fail. As the cavitation number in the pump decreases below the critical point, the pump fails to operate. There is concern regarding the behavior of the turbopump assembly arising from pump failure due to cavitation. It is necessary to verify the problems that may occur if the turbopump assembly operates under extreme conditions, such like the critical cavitation. This study performed tests to investigate the breakdown of pumps in the turbopump assembly. Tests were conducted with liquid nitrogen, water, and high-pressure air instead of the mediums used during actual operation of liquid oxygen, kerosene, and hot gas. The turbopump was tested at the design point of 27,000 rpm, while the inlet pressure of each pump was controlled to approach the critical cavitation number. The turbine power output was maintained during the tests. The results show that the breakdown point of the oxidizer pump using liquid nitrogen, which is a cryogenic medium, occurred at a lower cavitation number than during an individual component suction performance test using water. The fuel pump using water, meanwhile, experiences breakdown at similar cavitation numbers in both tests. As the breakdown of the pump occurs, the power required by that pump decreases, and the rotational speed of the turbopump increases. Compared with individual pump suction performance tests, this breakdown test can be used to determine the limit of the propellant inlet pressure of the turbopump and to characterize the behavior of the turbopump assembly when a breakdown occurs. Vibrations were also analyzed for tests at a high cavitation number and at the critical cavitation number. The vibration increased with breakdown and notable frequencies were analyzed.

Purge air is injected at the hub and shroud of axial turbines in order to avoid hot gas entering the gaps between stationary and rotating blade rows. The purge flows considerably interact with the main flow and influence the secondary flow like the tip leakage vortex. Therefore, at Graz University of Technology the flow in a product-representative one-and-a-half stage test turbine under the influence of purge flows was investigated. Four individual purge mass flows differing in flow rate, pressure, and temperature were injected through hub and tip cavities before and after the unshrouded high-pressure turbine rotor. In order to get more insight into the cavity flows and the flow evolution in the rotor this configuration is studied with a steady CFD simulation with and without purge flows. It was found that the secondary flow and especially the tip leakage vortex is significantly influenced by the purge flow which varies in circumferential direction. The differences between purge and zero-purge flow conditions are discussed with the help of radial distributions and contour plots of stream-wise vorticity. Streamlines allow to follow the path of the purge flows in the rotor and show the radial displacement of the secondary flow vortices. Wall streamlines describe the changes in the boundary layer flow and their effect on the vorticity after the trailing edge.

A centrifugal compressor requires a wide operating range as well as a high efficiency. At high pressure ratios, the impeller discharge velocity becomes transonic and effective pressure recovery in a vaned or vaneless diffuser is necessary. At high pressure ratios, a vaned diffuser is used as it has high pressure recovery, but may have a narrow operating range. At low flow, diffuser stall may trigger surge. At high flow, choking in the throat of the vanes may limit the maximum flow rate. A low solidity diffuser allows a good pressure recovery because it has vanes to guide the flow and a wide operating range as there is no geometrical throat to limit the maximum flow. In experimental studies at a pressure ratio around 4:1, the author has replaced vaned diffusers with a range of low solidity diffusers to try to broaden the operating range. The test results showed that the low solidity diffuser also chokes. In this paper, a virtual throat is defined and its existence is confirmed by flow visualization and pressure measurements. A method to select low solidity diffusers is proposed based on test data and the fundamental nature of the flow. The extension of the proposed method to the selection of a vaneless diffuser is examined and a design approach for a vaneless diffuser system to minimize surge flow rate without limiting the attainable maximum flow rate is proposed.

Gas turbine performance is highly dependent on the quality of the manufactured parts. Manufacturing variations in the parts can significantly alter the performance, especially efficiency and thus SFC. The legacy process is to accept variations within predefined profile tolerance limits and a few other qualitative parameters, mostly at a few, key two-dimensional aerofoil sections. With the widespread use of White light scans and other similar three-dimensional scans, this has improved to include the three-dimensional profile. The future however may lie with performance based quality assessment of manufactured parts, combined with quantitative surface quality assessment to implement an intelligent screening process for the parts. The adjoint method, typically used for shape optimization is adapted to provide a prediction of the impact on performance due to manufacturing variations. The work presented outlines a three stage quality assessment process for manufactured parts, involving three-dimensional profile tolerance based screening, followed by a surface curvature based screening and finally an Adjoint based performance prediction.

In an axial flow turbine, almost one-third of the total losses are caused by secondary flows, and the non-axisymmetric endwall profiling has been a major tool for years to reduce the secondary flow loss. This paper presents the non-axisymmetric endwall profile construction and optimization both for the stator hub and shroud of a high pressure turbine in the presence of an axisymmetric rotor. The flow simulation in the turbine was conducted by using steady RANS. The perturbation law of non-axisymmetric endwall was based on Bezier curves, and the commercially available optimization software NUMECA Fine/Design 3D was used to design the non-axisymmetric endwall. A genetic algorithm based on the artificial neural network was used as the optimization method. The objective function was aimed at maximizing the stage isentropic efficiency. The change in mass flow rate was kept less than 0.5% (relative) so that efficiency might not be influenced by the mass flow through the variation of the throat area. From the design point of view, the stator hub endwall was optimized at design conditions firstly, but the shroud endwall was kept constant, which resulted in an increase of stage efficiency because flow angles at stator exit were changed. The flow structures in the passage of stator were compared pre and post optimization by using 3-D streamlines in the vicinity of the endwall. Subsequently, the shroud endwall was optimized using the optimized non-axisymmetric hub as initial design. Due to hub and shroud endwall perturbation, the cross passage gradient and entropy were reduced, and the turbine stage efficiency at design conditions was calculated and the improvement in the efficiency was noticed. In addition, the improved hub and shroud contour were considered for off-design conditions as well, and efficiency was even more increased over a considerable off-design regime than at the design point condition.

The role of gas-to-wall temperature ratio on bypass transition along a highly loaded turbine guide vane is investigated through time-resolved heat flux measurements for different flow conditions. The tests are conducted in the von Karman Institute (VKI) isentropic compression tube (CT-2) facility. High response thin films mounted on the vane (VKI LS89 airfoil) coupled with analog circuits are used for the heat flux measurements. The first detectable wall heat flux fluctuations denote the onset of transition which is also evaluated by both the heat transfer coefficient and the intermittency factor distributions along the suction side. The exit Mach number is kept constant during each test by means of a downstream sonic throat while the gas-to-wall temperature ratio is varied from 1.14 to 1.51 by changing the inlet gas temperature. Four test cases are considered for different exit Mach numbers (0.52 and 0.86) and freestream turbulence intensity levels (0.8% and 5.3%) while the isentropic exit chord Reynolds number is maintained at 10 6 . In the present test campaign, the length and the evolution of the phenomenon indicate a measurable effect of the gas-to-wall temperature ratio for test cases with mild pressure gradients.

The subject of the paper is a high speed unshrouded centrifugal compressor in which rotating instabilities have been measured near the stage stall point. The impeller is studied numerically by means of unsteady RANS simulations, and the results are compared to experimental measurements. Instead of rotating instabilities, the numerical results directly capture a rotating stall pattern in which a tornado-like separation vortex is shed due to a separation at the impeller leading edge, and propagates around the circumference. The vortex has one end attached to the casing and the other end attached to the pressure side of the blade. Its propagation speed is within 10% from the measured speed of the rotating instabilities. Because of the high pressure gradient the tip leakage flow crosses the impeller front and thus convects the vortex in front of the adjacent blade. The spillage of the vortex in the adjacent channel convects radial and azimuthal vorticity onto the next blade. This triggers the inception of a new vortex, and induces the propagation of the rotating stall cells.

Flow exiting the combustor is highly turbulent and contains significant spatial gradients of pressure and temperature. The high pressure turbine nozzle vanes operating in this environment redistribute these spatial gradients and impact the inflow characteristics of the turbine rotor blades. The present study investigates the redistribution of total temperature through a turbine nozzle vane. Numerical investigation was performed using three-dimensional RANS analysis. Simulations were conducted using the Wilcox k–ω turbulence model and Shear Stress Transport (SST) with and without γ–Re θ transition model. Experimental measurements were obtained in an annular nozzle cascade facility. Two sets of inlet conditions were considered. The first was a nominally uniform total temperature. The second had a span-wise variation of total temperature. Both sets of inlet conditions had nominally the same inlet total pressure and inlet Mach number. Span-wise redistribution was evaluated using the circum-ferentially averaged total temperature profile at a plane downstream of the nozzle. Physical arguments about the influence of nozzle secondary flows on this redistribution are presented.