Continuous reduction of fuel consumption for a wide range of operating conditions leads to a high efficiency demand for all engine parts of modern jet engines and especially the compressor. To meet these requirements a two-part Variable Inlet Guide Vane (VIGV), composed of a fixed strut and a variable flap, can be used. Besides the aerodynamic aspects, the VIGV strut is a substantial part for the structural integrity of the compressor. The aerodynamic design optimization of such a VIGV, located upstream of the first rotor of a 2.5 stage low pressure compressor, under the conditions of three different operating points is presented in this paper. In a previous study the shape of the axial gap between strut and flap was optimized without changing the envelope of both parts [1]. The new design tool SplitBlade, developed at the DLR, enables the creation of an axial gap and has been integrated in the design process of the in-house optimization tool AutoOpti. The target of the optimization was to decrease the total pressure loss coefficients for all three operating points. The design optimization presented in this paper is more complex by allowing the VIGV blade geometry to change. The basic dimensions of the VIGV such as the axial chord and the maximum profile thickness are still frozen. In total, 88 parameters are free to change in the optimization process. Additionally to the main target of loss reduction, the circumferential outflow angles are restricted to maintain the deflection of the blade and therewith the required rotor inflow conditions to ensure the operability of the entire compressor in the whole working range. The final result is a two-part VIGV with an axial gap, which is optimized in terms of total pressure losses in three operating points. Compared to a reference geometry without an axial gap, the losses are almost equal at nominal speed, and about one to two percentage points higher in the two part speed operating points.

As bypass-ratio in modern aero engines is continuously increasing over the last decades, the radial offset between low pressure compressor (LPC) and high pressure compressor (HPC), which needs to be overcome by the connecting s-shaped intermediate compressor duct (ICD), is getting higher. Due to performance and weight saving aspects the design of shorter and therefore more aggressive ducts has become an important research topic. In this paper an already aggressive design (with respect to current aero engines) of an ICD with integrated outlet guide vane (OGV) is used as a baseline for an aerodynamic optimization. The aim is to shorten the duct even further while maintaining it separation free. The optimization is broken down into two steps. In the first optimization-step the baseline design is shortened to a feasible extent while keeping weak aerodynamic restrictions. The resulting highly aggressive duct (intermediate design), which is shortened by 19 % in axial length with respect to the baseline, shows separation tendencies of low momentum fluid in the strut/hub region. For the second step, the length of the optimized duct design is frozen. By implementing new design features in the process of the optimizer, this optimization-step aims to eliminate separation and to reduce separation tendencies caused by the aggressive shortening. In particular, these features are: a nonaxisymmetric endwall contouring and parametrization of the strut and the OGV to allow for changes in lift and turning in both blade designs. By comparison of the three designs: Baseline, intermediate (separating flow) and final design, it can be shown, that it is possible to decrease length of the already aggressive baseline design even further, when adding a nonaxisymmetric endwall contouring and changes in blade shape of the strut and OGV. Flow separation can be eliminated while losses are kept low. With a more aggressive and therefore shorter duct the engine length and weight can be reduced. This in turn leads to lighter aircrafts, less fuel consumption and lower CO 2 and NO x emissions.

This paper deals with the reduction of aerodynamically generated noise in passenger car Cooling-Fan-Modules (CFM), caused by the interaction between the impeller and the downstream-located strut configuration of the axial fan. Even after the car engine is switched off, the fan remains active, as long as cooling is required for certain vehicle components. Especially after a car has been parked in closed parking areas, in close proximity to residential buildings or public places, the noise emission can be a problem. This issue is addressed by dampening the rotor-stator-interaction through passive construction measures. In order to ensure optimal noise reduction, 8 critical design features of the struts are identified and investigated using statistical design of experiment methods (DoE). Based on the results, dedicated insights about the effects of concrete strut features on significant regions of the acoustic fan spectrum are obtained. Furthermore, an optimized strut configuration is derived and metrologically validated using a polyoptimization method. Compared to a current serial baseline configuration, a reduction of the overall sound pressure level by 2.6 dB(A), as well as a reduction of the blade passage frequency tone by 17.6 dB(A) is achieved.

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 efficacy of integrating the lobed mixer with the core flow deswirling struts to create a single component for reducing the exhaust system length, beyond that attainable through mixer optimization alone, has been investigated. This investigation has been conducted via CFD simulations of a medium-bypass turbofan exhaust system at engine cruise representative conditions. Comparative analysis shows that integration augmented thrust output by about 0.02% while total pressure loss was increased by 3.6%. The aim of the study, to show that this new integrated design would have either minimal impact on or improve exhaust system performance, was confirmed. Comparisons of the flow fields and characteristic quantities downstream of the mixer also showed minimal impact on flow through the nozzle. The deswirling strut was offset by 0.65 D h axially when integrated with the mixer, therefore it can be concluded that the exhaust system ducting could be reduced in length by this same measure — saving engine weight in the process.

A simulation of the flow through a four-stage low pressure turbine (LPT) of an aero-engine is performed using an eddy-viscosity based RANS model. An analysis of secondary flow details in the vicinity of endwalls and strut surface is performed. An interaction of wakes generated by the swirler vanes and end-wall boundary layers is also studied. The CFD results are compared with own test rig data of LPT obtained in Polonia Aero Lab in Zielonka (Poland). Global flow parameters were reproduced in good agreement with experiment (difference less than 1.5%). Satisfactory agreement between measured and predicted pressure distribution on the surface of inlet strut has been obtained.

The protection of the inlet components of an aircraft engine from the adverse effects of ice accretion has been a crucial design problem since the very early years of flight. Therefore, the configuration with efficient heat transfer is the focus on the design of a hot-air anti-icing system in an aeroengine. This paper studies experimentally on the heat transfer in a convergent channel with pin fin within the strut. Experiments are carried out by using a transient liquid crystal technique. The dimensionless lateral distance ( S/d , where S is the distance between the leading edge and pin fin, d is the diameter of the impingement hole), the dimensionless vertical distance ( S1/d , where S1 is the distance between pin fins) as well as the pitch ratio ( D/d , where D is the diameter of the pin fin) of the pin fin and impingement hole are respectively investigated to study the heat transfer on the surface of convergent channel. The Reynolds number based on the hydraulic diameter of the impingement hole ranges from 6300 to 12700. Within the experimental range, the result shows that the Average Nusselt Number increases at first and then decreases as the dimensionless lateral distance S/d increases. The Average Nusselt Number has the highest value when S/d = 2.5;When the dimensionless vertical distance S1/d increases, the Average Nusselt Number goes up at first and then reduces .The Average Nusselt Number has the highest value when S1/d = 3 ;The increasing pitch ratio D/d leads to a lower Average Nusselt Number. The Average Nusselt Number has the highest value when D/d = 1/3 . The Average Nusselt Number increases as the Reynolds number increases.

The influence of a cylindrical strut shortly downstream of the bladerow on the vibration behavior of the last stage rotor blades of a single stage LP model steam turbine was investigated in the present study. Steam turbine retrofits often result in an increase of turbine size, aiming for more power and higher efficiency. As the existing LP steam turbine exhaust hoods are generally not modified, the last stage rotor blades frequently move closer to installations within the exhaust hood. To capture the influence of such an installation on the flow field characteristics, extensive flow field measurements using pneumatic probes were conducted at the turbine outlet plane. In addition, time-resolved pressure measurements along the casing contour of the diffuser and on the surface of the cylinder were made, aiming for the identification of pressure fluctuations induced by the flow around the installation. Blade vibration behavior was measured at three different operating conditions by means of a tip timing system. Despite the considerable changes in the flow field and its frequency content, no significant impact on blade vibration amplitudes were observed for the investigated case and considered operating conditions. Nevertheless, time-resolved pressure measurements suggest that notable pressure oscillations induced by the vortex shedding can reach the upstream bladerow.

This paper describes studies completed using a quarter-scaled rig to assess the impact of turbine exit swirl angle and strut stagger on a turbine exhaust system consisting of an integral diffuser-collector. Advanced testing methods were applied to ascertain exhaust performance for a range of inlet conditions aerodynamically matched to flow exiting an industrial gas turbine. Flow visualization techniques along with complementary Computational Fluid Dynamics (CFD) predictions were used to study flow behavior along the diffuser endwalls. Complimentary CFD analysis was also completed with the aim to ascertain the performance prediction capability of modern day analytical tools for design phase and off-design analysis. The K-Epsilon model adequately captured the relevant flow features within both the diffuser and collector, and the model accurately predicted the recovery at design conditions. At off-design conditions, the recovery predictions were found to be pessimistic. The integral diffuser-collector exhaust accommodated a significant amount of inlet swirl without a degradation in performance, so long as the inlet flow direction did not significantly deviate from the strut stagger angle. Strut incidence at the hub was directly correlated with reduction in overall performance, whereas the diffuser-collector performance was not significantly impacted by strut incidence at the shroud.

This paper deals with the influence of high-pressure turbine purge flows on the aerodynamic performance of turbine center frames. Measurements were carried out in a product-representative one and a half stage turbine test setup, installed in the Transonic Test Turbine Facility at Graz University of Technology. The rig allows testing at engine-relevant flow conditions, matching Mach, Reynolds, and Strouhal number at the inlet of the turbine center frame. Four individual purge mass flows differing in flow rate, pressure, and temperature were injected through the hub and tip, forward and aft cavities of the unshrouded high-pressure turbine rotor. Two turbine center frame designs (differing in area distribution and inlet-to-exit radial offset), equipped with non-turning struts, were tested and compared. For both configurations, aerodynamic measurements at the duct inlet and outlet as well as oil flow visualizations through the turbine center frame were performed. The acquired measurement data illustrate that the interaction of the ejected purge flow with the main flow enhances the secondary flow structures through the turbine center frame duct. Depending on the purge flow rates, the radial migration of purge air onto the strut surfaces directly impacts the loss behavior of the duct. While the duct loss is demonstrated to be primarily driven by the core flow between two duct struts, the losses associated with the flow close to the struts and in the strut wakes are highly dependent on the relative position between the high-pressure turbine vane and the strut leading edge, as well as the interaction between vane wake and ejected purge flow. Hence, while the turbine center frame duct pressure loss depends on the duct geometric characteristics it is also influenced by the presence and rate of the high-pressure turbine purge flows. This first-time experimental assessment demonstrates that a reduction in the high-pressure turbine purge and cooling air requirement not only benefits the engine system performance by decreasing the secondary flow taken from the high-pressure compressor but also by lowering the turbine center frame total pressure loss.

Decreasing drag on aircraft components was beneficial towards improving fuel economy and operational range. A generic axisymmetric nacelle-strut configuration typical of those housing fuselage-mounted engines was evaluated at a Reynolds number of 6 × 10 5 based on the nacelle maximum diameter d = 26.5 cm and an angle of attack of 20 deg. It was estimated that drag could be reduced by 20%. Three case studies were evaluated that added a fillet to the nacelle-strut corner, vortex generating triangular tabs, and flow-path obstructing vanes to improve flow control by reducing suction-side separation. Experimental results showed that a 0.11 d radius of curvature fillet reduced drag by 8% with respect to the baseline case. Numerical results employing the realizable k-ε turbulence model with wall functions predicted no improvements with the tabs and an 8% reduction with the vanes.

The flow leaving the high pressure turbine should be guided to the low pressure turbine by an annular diffuser, which is called as the intermediate turbine duct. Flow separation, which would result in secondary flow and cause great flow loss, is easily induced by the negative pressure gradient inside the duct. And such non-uniform flow field would also affect the inlet conditions of the low pressure turbine, resulting in efficiency reduction of low pressure turbine. Highly efficient intermediate turbine duct cannot be designed without considering the effects of the rotating row of the high pressure turbine. A typical turbine model is simulated by commercial computational fluid dynamics method. This model is used to validate the accuracy and reliability of the selected numerical method by comparing the numerical results with the experimental results. An intermediate turbine duct with eight struts has been designed initially downstream of an existing high pressure turbine. On the basis of the original design, the main purpose of this paper is to reduce the net aerodynamic load on the strut surface and thus minimize the overall duct loss. Full three-dimensional inverse method is applied to the redesign of the struts. It is revealed that the duct with new struts after inverse design has an improved performance as compared with the original one.

Three different optimization methods using genetic algorithms have been developed, aiming to achieve better aggressive intermediate turbine duct (ITD) performance. To overcome defects of simple genetic algorithms, a niche genetic algorithm is used, for its better adaptability to multi-peak function. These three methods are two-dimensional optimization for pursuing the highest static pressure coefficient; two-dimensional optimization via controlling static pressure coefficient; and a further three-dimensional optimization. The second method introduces a restrainer to make sure the maximum value of static pressure coefficient gradient less than a limitation. A simulation case, an ITD with eight struts, was implemented to demonstrate the capabilities of the presented optimization methods. Compared to the baseline ITD, the results show as follow. The first method obtains a best static pressure coefficient but a severe separation. The second optimization method with static pressure coefficient gradient control can definitely suppress separation in the ITD, the second method also obtains better static pressure coefficient and the lowest total pressure loss coefficient. The third method, a further three-dimensional optimization can obtain better ITD overall performance because of a more realistic simulation. Nevertheless, the second two-dimensional optimization method can get good enough results while it is apparently much more time-saving compared to three-dimensional one, which make it more suitable for engineering applications.

A detailed numerical simulation is presented to investigate the new de-swirling methods and their effect on the mixing mechanisms of a turbofan mixer with 12 lobes. The numerical simulation employed a commercial solver, ANSYS CFX, using k-ω SST model. The core-to-bypass temperature ratio and pressure ratio were set to 2.59, and 0.97 respectively, giving the Mach number of 0.66 and bypass ratio of 2.65 at mixing nozzle outlet. The inlet swirl typically accelerates the jet-flow mixing by enhancing the vortices intensity and interaction, but leakage swirling flow can cause a three-dimensional separation bubble and the recirculation zone resulting in the dramatic increasing the total pressure loss and thrust loss. Removal of the leakage swirling flow between the lobes’ trough and centre-body was the key to limit the negative influence of inlet swirl. Two IGV design were investigated, DS1 and DS2. DS1 was installed at the upstream of the lobed mixer, could remove the negative effect of inlet swirl properly, but also inhibited the active role of the inlet swirl. The total pressure and thrust loss reduced by 0.31% and 3.8%, respectively, but the mixing efficiency also decreased by 1.72%. DS2, an integrated strut with the lobed mixer design, not only ensured the structure strength of the lobed mixer, but also reduced the length and weight of the exhaust system. This method suppressed the flow separation bubble on centre-body to some extent, and eliminated the recirculation zone downstream of the cenrebody, resulting in the total pressure loss decrease of 0.31% and thrust gain of 3.63%. On the other hand, the method DS2 also made full use of the inlet swirl to enhance the jet-flow mixing, resulting in the mixing efficiency increased 1.54% compared with that of the DS1 case. Under the off-design conditions with the incidence angle of ±10°, the aerodynamic performance of the DS2 cases didn’t changed too much such as the DS1 cases.

The jetted hot film could affect the trajectories of the water droplets near the aero-engine inlet strut surface, which equipped with the ice protection system combined the internal impingement heat transfer and the external hot air film heating. To evaluate the droplets impingement characteristics of four ice-protection structures designed with different film-slot jet angles, a droplets impingement computation method based on Eulerian framework was developed and validated. The influences of film-slot angle and blowing ratio on the impingement characteristics for droplet diameter of 20μm were investigated and the jet vortex was found to be an important factor. The results indicated that the local collection efficiency and the impingement limitation could decrease significantly due to the blowing from the external hot-film, and the influence would be more significant in case that the film-slot was closer to the leading edge. For example, the average local collection efficiencies of four typical configurations with different slot angles and positions decreased 82%, 8%, 1% and 0.5% respectively comparing to those without air film. Besides that, the maximum local collection efficiency and the impingement limitation decreased with the increasing blowing ratios, and the film-slot nearest to the leading edge was most sensitive to the blowing ratio. It was also found that no droplets impinged on the rear surface after the jet slots at some higher blowing ratios.

In order to achieve high working efficiency, modern gas turbines operate at high temperature which is close to the melting points of metal alloys. However, the support of turbine end suffers the thermal deformation. And the journal center position is also changed due to the effects of high temperature and shaft gravity. Tangential or radial supporting structures, which are composed of supporting struts, diffuser cones, hot and cooling fluid channel, are widely used in gas turbine hot end. Cooling technology is usually used to keep the bearing temperature in a reasonable range to meet requirements of strength and deformation of the supporting struts. In this paper, three major assumptions are proposed: (a) radiation is not considered, (b) cooling flow system is only partially modeled and analysis assumes significantly higher cooling flow that is not typical for current engines, and (c) only steady state heat transfer is considered. And a 3D fluid-solid coupled model based on finite-element method (FEM) is built to analyze the performances of both the tangential and the radial support. The temperature distribution, thermal deformation and stress of supports are obtained from CFD and strength analysis. The results show that either the tangential or radial support is used in a 270MW gas turbine; the thermal stress is about 90.3% of total stress which is produced by both thermal effects and shaft gravity. Comparing to the results from radial supports, it can be seen that the struts stress and position variation of journal center of tangential support are smaller. Due to a rotational effect of the bearing housing caused by the deformation of the tangential struts, the thermal stress in these tangential struts can be relieved to some extent. When both thermal effect and shaft gravity are considered, the stress of each tangential supporting strut is almost uniformly distributed, which is beneficial to the stability of rotor system in the gas turbine.

A comprehensive experimental investigation was initiated to evaluate the aerodynamic performance of a gas turbine exhaust diffuser/collector for various strut geometries over a range of inlet angle. The test was conducted on a 1/12th scale rig developed for rapid and accurate evaluation of multiple test configurations. The facility was designed to run continuously at an inlet Mach number of 0.40 and an inlet hydraulic diameter-based Reynolds number of 3.4×10 5 . Multi-hole pneumatic pressure probes and surface oil flow visualization were deployed to ascertain the effects of inlet flow angle and strut geometry. Initial baseline diffuser-only tests with struts omitted showed a weakly increasing trend in pressure recovery with increasing swirl, peaking at 14° before rapidly dropping. Tests on profiled struts showed a similar trend with reduced recovery across the range of swirl and increased recovery drop beyond the peak. Subsequent tests for a full diffuser/collector configuration with profiled struts revealed a rising trend at lower swirl when compared to diffuser-only results, albeit with a reduction in recovery. When tested without struts, the addition of the collector to the diffuser not only reduced the pressure recovery at all angles but also resulted in a shift of the overall characteristic to a peak recovery at a lower value of swirl. The increased operation range associated with the implementation of struts in the full configuration is attributed to the de-swirling effects of the profiled struts. In this case the decreased swirl reduces the flow asymmetry responsible for the reduction in pressure recovery attributed to the formation of a localized reverse-flow vortex near the bottom of the collector. This research indicates that strut setting angle and, to a lesser extent, strut shape can be optimized to provide peak engine performance over a wide range of operation.

The flow and heat transfer inside a short airfoil shaped turbine exhaust strut was studied experimentally and computationally. A combination of transient and steady IR thermographic techniques were used to measure the heat transfer, and showed that the streamwise evolution of the heat transfer could be reliably described using a power law type relationship. Steady RANS CFD using the realizable k-epsilon was unable to accurately predict the observed heat transfer, with the main limitations of the model being a) a fundamental inability to calculate secondary flows and b) an underestimation of the restriction on boundary layer growth rates imposed by the opposing walls. The experiments showed that heat transfer rates were much higher in thinner regions of the strut, and lower in thicker regions. The CFD predicted an essentially uniform heat transfer rate, independent of the local strut thickness.

This paper deals with the investigation on the acoustics of different turning mid turbine frames (TMTF) in the two-stage two-spool test turbine located at the Institute for Thermal Turbomachinery and Machine Dynamics (ITTM) of Graz University of Technology. The facility is a continuously operating cold-flow open-circuit plant which is driven by pressurized air. The flow path consists of a transonic turbine stage (HP) followed by a low pressure turbine stage made of a turning mid turbine frame (TMTF) and a counter-rotating low pressure rotor. Downstream of the low pressure turbine a measurement section is instrumented with acoustic sensors. Three TMTF setups have been investigated at engine like flow conditions. The first configuration (C1) consists of 16 highly 3D-shaped turning struts. The goal of the second design (C2) was to reduce the length of the TMTF by 10% without increasing the losses and providing comparable inflow to the LP turbine rotor. This was achieved by applying 3D-contoured endwalls at the hub. The third one (C3) is a new embedded concept for the turning mid turbine frame with two zero-lift splitters placed into the strut passages. In total 48 vanes (16 struts plus 32 splitter vanes) guide the flow from the HP rotor to the LP rotor. The comparison in terms of noise generation and propagation of the turbines is done by the microphones signal spectra, the emitted sound pressure and sound power level of each TMTF setup. Therefore the acoustic field is characterized by azimuthal and radial modes by means of a microphone array at the outer casing traversed over 360 degrees. By comparing the first two setups (C1 and C2) in terms of noise generation the propagating modes due to the HP turbine were found to be the same, while a difference of 5 dB in amplitude of the modes related to the LP turbine was found due to the different axial spacing between both setups. In the multi-splitter configuration (C3), the overall sound power level depending on the blade passing frequency (BPF) of the HP turbine is reduced by 7 dB and depending on the BPF of the LP turbine by 4 dB compared to C1, respectively. The overall effect is a reduction of the noise emission for the HP turbine due to the embedded design as well as for the LP turbine due to increasing the axial spacing between the TMTF Vanes and LP Blades on the one hand and considerably due to the embedded design on the other hand.

The trend for higher overall pressure ratios means that turbine entry temperatures are continually increasing. Furthermore, the development of lean, low-emission combustion systems reduces the availability of cooling air and is accompanied by new problems at the combustor/turbine interface. For example, the exit temperature traverse differs from that found in traditional rich-burn combustors with increased swirl and a much flatter profile. Effectively cooling the turbine components is becoming increasingly difficult. One solution is to employ cooled cooling air (CCA) where some of the compressor efflux is diverted for additional cooling in a heat exchanger located in the by-pass duct. An example CCA system is presented which includes an off-take within the dump cavity and the addition of radial struts within the pre-diffuser through which the cooled air is returned to the engine core. This paper addresses the impact this CCA system has on the combustion system external aerodynamics. This included the development of a fully annular, isothermal test facility which incorporated a bespoke 1.5 stage axial compressor, engine relevant outlet guide vanes, pre-diffuser and combustor geometry. A datum aerodynamic performance was established for a non-CCA configuration with a clean, un-strutted pre-diffuser. Results for this baseline CCA system demonstrated that inclusion of a bleed in the dump cavity had limited effect on the overall flow field. However, the inclusion of struts within the pre-diffuser caused a reduction in area ratio and a notable increase in system loss. Consequently an alternative pre-diffuser was designed (using CFD) with the aim of increasing the area ratio back to that of the un-strutted datum. A so-called hybrid diffuser was designed in which the CCA bleed was moved to the pre-diffuser outer wall. The bleed was then used to re-energize the boundary layer, preventing flow separation, enabling the area ratio to be increased close to the datum value. The mechanisms of the hybrid diffuser are complex; the geometry of the off-take and its location with respect to the OGV and strut leading edge were seen to be critical. Experimental evaluation of the final design demonstrated the effective operation the hybrid diffuser with the result that the system loss returned to a level close to that of the datum. Only small differences were seen in the overall flow field.