The aerodynamic interaction of upstream and downstream blade rows can have a significant impact on the forced response of the compressor. Previously, the authors carried out the forced response analysis of a three-row stator-rotor-stator (S1-R2-S2) configuration from a 3.5-stage compressor. However, since the stator vane counts in both the stators (S1 and S2) were the same, it was not possible to separate the excitations from both the rows as they excited the rotor at the same frequency. Hence, a new configuration was developed and tested in which the stator 1 blade count was changed to 38 and stator 2 blade count was maintained at 44 in order to study the individual influences of the stator on the embedded rotor. By using this method, the excitations from both rows can be determined, and the excitations can be quantified to determine the row having the maximum influence on the overall forcing. To achieve this, two sets of simulations were carried out. The three-row stator-rotor (S1-R2-S2) simulation was carried out at both the 38EO (engine order) and 44EO crossings at the peak efficiency (PE) operating condition. The two-row stator-rotor analysis (S1-R2) was carried out at the 38EO crossing, and the other two-Row (R2-S2) analyses were carried out at the 44EO crossing. The steady aerodynamics was preserved in both the cases. A study was done to determine the contribution of wave reflections from the stator inlet and exit planes to the forcing function. Two conclusions drawn from this study are as follows: (1) the modal force value decreased after the upstream stator was removed, which proved that wave reflections from this stator were significant and (2) the increase in modal force was in-line with experimental observations.

As aero gas turbine designs strive for ever greater efficiencies, the trend is for engine overall pressure ratios to rise. Although this provides greater thermal efficiency, it means that cycle temperatures also increase. One potential solution to managing the increasing temperatures is to employ a cooled cooling air system. In such a system, a purge flow into the main gas path downstream of the compressor will be required to prevent hot gas being ingested into the rotor drive cone cavity. However, the main gas path in compressors is aerodynamically sensitive and it is important to understand, and mitigate, the impact such a flow may have on the compressor outlet guide vanes, pre-diffuser, and the downstream combustion system aerodynamics. Initial computational fluid dynamics (CFD) predictions demonstrated the potential of the purge flow to negatively affect the outlet guide vanes and alter the inlet conditions to the combustion system. The purge flow modified the incidence onto the outlet guide vane, at the hub, such that the secondary flows increased in magnitude. An experimental assessment carried out using an existing fully annular, isothermal test facility confirmed the CFD results and importantly demonstrated that the degradation in the combustor inlet flow resulted in an increased combustion system loss. At the proposed purge flow rate, equal to ∼1% of the mainstream flow, these effects were small with the system loss increasing by ∼4%. However, at higher purge flow rates (up to 3%), these effects became notable and the outlet guide vane and pre-diffuser flow degraded significantly with a resultant increase in the combustion system loss of ∼13%. To mitigate these effects, CFD was used to examine the effect of varying the purge flow swirl fraction in order to better align the flow at the hub of the outlet guide vane. With a swirl fraction of 0.65 (x rotor speed), the secondary flows were reduced below that of the datum case (with no purge flow). Experimental data showed good agreement with the predicted flow topology and performance trends but the measured data showed smaller absolute changes. Differences in system loss were measured with savings of around 10% at the turbine feed ports for a mass flow ratio of 1% and a swirl fraction of 0.65.

In this paper, the effect of aerodynamic mistuning on stability of a compressor cascade is studied. The experiments have been carried out at a low-speed test facility of the Technische Universität Berlin. The test section contains a linear cascade with compressor blades that are forced to oscillate in sinusoidal pitching motion. The aerodynamic mistuning is realized by a blade-to-blade stagger angle variation, and three mistuning patterns have been investigated: one-blade mis-staggering, alternating mis-staggering, and random mis-staggering. Mis-staggering can have a stabilizing or destsabilizing effect, but depends strongly on the amount of detuning that alters the flow passage. For positive stagger angle variation for the one-blade and alternating mis-staggering, the trend of the damping curve was maintained, in the sense that the unstable interblade phase angles (IBPAs) remained unstable. For negative stagger angle variation, one IBPA shifted from stable to unstable. For the random pattern, only very moderate changes are observed. The cascade stability was not noticeably affected by the aerodynamic mistuning.

Variable pitch fans are of interest for future low-pressure ratio fan systems since they provide improved operability relative to fixed pitch fans. If they can also be re-pitched such that they generate sufficient reverse thrust they could eliminate the engine drag and weight penalty associated with bypass duct thrust reversers. This paper sets out to understand the details of the 3D fan stage flow field in reverse thrust operation. This study uses the Advanced Ducted Propulsor variable pitch fan test case, which has a design fan pressure ratio of 1.29. Comparison with spanwise probe measurements show that the computational approach is valid for examining the variation of loss and work in the rotor in forward thrust. The method is then extended to a reverse thrust configuration using an extended domain and appropriate boundary conditions. Computations, run at two rotor stagger settings, show that the spanwise variation in relative flow angle onto the rotor aligns poorly to the rotor inlet metal angle. This leads to two dominant rotor loss sources: one at the tip associated with positive incidence and the second caused by negative incidence at lower span fractions. The second loss is reduced by opening the rotor stagger setting, and the first increases with rotor suction surface Mach number. The higher mass flow at more open rotor settings provide higher gross thrust, up to 49% of the forward take-off value, but is limited by the increased loss at high speed.

This paper describes the development and initial application of an adjoint harmonic balance (HB) solver. The HB method is a numerical method formulated in the frequency domain which is particularly suitable for the simulation of periodic unsteady flow phenomena in turbomachinery. Successful applications of this method include unsteady aerodynamics as well as aeroacoustics and aeroelasticity. Here, we focus on forced response due to the interaction of neighboring blade rows. In the simulation-based design and optimization of turbomachinery components, it is often helpful to be able to compute not only the objective values—e.g., performance data of a component—themselves but also their sensitivities with respect to variations of the geometry. An efficient way to compute such sensitivities for a large number of geometric changes is the application of the adjoint method. While this is frequently used in the context of steady computational fluid dynamics (CFD), it becomes prohibitively expensive for unsteady simulations in the time domain. For unsteady methods in the frequency domain, the use of adjoint solvers is feasible but still challenging. The present approach employs the reverse mode of algorithmic differentiation (AD) to construct a discrete adjoint of an existing HB solver in the framework of an industrially applied CFD code. The paper discusses implementational issues as well as the performance of the adjoint solver, in particular regarding memory requirements. The presented method is applied to compute the sensitivities of aeroelastic objectives with respect to geometric changes in a turbine stage.

Objective of this paper is to analyze the consequences of borescope blending repairs on the aeroelastic behavior of a modern high pressure compressor (HPC) blisk. To investigate the blending consequences in terms of aerodynamic damping and forcing changes, a generic blending of a rotor blade is modeled. Steady-state flow parameters like total pressure ratio, polytropic efficiency, and the loss coefficient are compared. Furthermore, aerodynamic damping is computed utilizing the aerodynamic influence coefficient (AIC) approach for both geometries. Results are confirmed by single passage flutter (SPF) simulations for specific interblade phase angles (IBPA) of interest. Finally, a unidirectional forced response analysis for the nominal and the blended rotor is conducted to determine the aerodynamic force exciting the blade motion. The frequency content as well as the forcing amplitudes is obtained from Fourier transformation of the forcing signal. As a result of the present analysis, the change of the blade vibration amplitude is computed.

This work, a continuation of a series of investigations on the aerodynamics of aggressive interturbine ducts (ITD), is aimed at providing detailed understanding of the flow physics and loss mechanisms in four different ITD geometries. A systematic experimental and computational study was carried out by varying duct outlet-to-inlet area ratios (ARs) and mean rise angles while keeping the duct length-to-inlet height ratio, Reynolds number, and inlet swirl constant in all four geometries. The flow structures within the ITDs were found to be dominated by the boundary layer separation and counter-rotating vortices in both the casing and hub regions. The duct mean rise angle determined the severity of adverse pressure gradient in the casing's first bend, whereas the duct AR mainly governed the second bend's static pressure rise. The combination of upstream wake flow and the first bend's adverse pressure gradient caused the boundary layer to separate and intensify the strength of counter-rotating vortices. At high mean rise angle, the separation became stronger at the casing's first bend and moved farther upstream. At high ARs, a two-dimensional separation appeared on the casing and resulted in increased loss. Pressure loss penalties increased significantly with increasing duct mean rise angle and AR.

During engine operation, fan casing abradable liners are worn by the blade tip, resulting in the formation of trenches. This paper describes the influence of these trenches on the fan blade tip aerodynamics. A detailed understanding of the fan tip flow features for cropped and trenched clearances is first developed. A parametric model is then used to model trenches in the casing above the blade tip and varying blade tip positions. It is shown that increasing clearance via a trench reduces performance by less than increasing clearance through cropping the blade tip. A response surface method is then used to generate a model that can predict fan efficiency for a given set of clearance and trench parameters. This model can be used to influence fan blade design and understand engine performance degradation in service. It is shown that an efficiency benefit can be achieved by increasing the amount of tip rubbing, leading to a greater portion of the tip clearance sat within the trench. It is shown that the efficiency sensitivity to clearance is biased toward the leading edge (LE) for cropped tips and the trailing edge (TE) for trenches.

Flow in an intermediate turbine duct (ITD) is highly complex, influenced by the upstream turbine stage flow structures, which include tip leakage flow and nonuniformities originating from the upstream high pressure turbine (HPT) vane and rotor. The complexity of the flow structures makes predicting them using numerical methods difficult, hence there exists a need for experimental validation. To evaluate the flow through an intermediate turbine duct including a turning vane, experiments were conducted in the Oxford Turbine Research Facility (OTRF). This is a short duration high speed test facility with a 3/4 engine-sized turbine, operating at the correct nondimensional parameters for aerodynamic and heat transfer measurements. The current configuration consists of a high pressure turbine stage and a downstream duct including a turning vane, for use in a counter-rotating turbine configuration. The facility has the ability to simulate low-NO x combustor swirl at the inlet to the turbine stage. This paper presents experimental aerodynamic results taken with three different turbine stage inlet conditions: a uniform inlet flow and two low-NO x swirl profiles (different clocking positions relative to the high pressure turbine vane). To further explain the flow through the 1.5 stage turbine, results from unsteady computational fluid dynamics (CFD) are included. The effect of varying the high pressure turbine vane inlet condition on the total pressure field through the 1.5 stage turbine, the intermediate turbine duct vane loading, and intermediate turbine duct exit condition are discussed and CFD results are compared with experimental data. The different inlet conditions are found to alter the flow exiting the high pressure turbine rotor. This is seen to have local effects on the intermediate turbine duct vane. With the current stator–stator vane count of 32-24, the effect of relative clocking between the two is found to have a larger effect on the aerodynamics in the intermediate turbine duct than the change in the high pressure turbine stage inlet condition. Given the severity of the low-NO x swirl profiles, this is perhaps surprising.

This paper presents the development and aerothermal investigation of the integrated combustor vane concept for power generation gas turbines with individual can combustors. In this novel concept, first introduced in 2010, the conventional nozzle guide vanes (NGVs) are removed and flow turning is achieved by vanes that extend the combustor walls. The concept was developed using the in-house computational fluid dyanamics (CFD) code TBLOCK. Aerothermal experiments were conducted using a modular high-speed linear cascade, designed to model the flow at the combustor–vane interface. The facility is comprised of two can combustor transition ducts and either four conventional vanes (CVs) or two integrated vanes (IVs). The experimental study validates the linear CFD simulations of the IV development. Annular full-stage CFD simulations, used to evaluate aerodynamics, heat transfer, and stage efficiency, confirm the trends of the linear numerical and experimental results, and thus demonstrate the concept's potential for real gas turbine applications. Results show a reduction of the total pressure loss coefficient at the exit of the stator vanes by more than 25% due to a reduction in profile and endwall loss. Combined with an improved rotor performance demonstrated by unsteady stage simulations, these aerodynamic benefits result in a gain in stage efficiency of above 1%. A distinct reduction in heat transfer coefficient (HTC) levels on vane surfaces, on the order of 25–50%, and endwalls is observed and attributed to an altered state of boundary layer (BL) thickness. The development of IV's endwall- and leading edge (LE)-cooling geometry shows a superior surface coverage of cooling effectiveness, and the cooling requirements for the first vane are expected to be halved. Moreover, by halving the number of vanes, simplifying the design and eliminating the need for vane LE film cooling, manufacturing and development costs can be significantly reduced.

The computation of the final, friction saturated limit cycle oscillation amplitude of an aerodynamically unstable bladed-disk in a realistic configuration is a formidable numerical task. In spite of the large numerical cost and complexity of the simulations, the output of the system is not that complex: it typically consists of an aeroelastically unstable traveling wave (TW), which oscillates at the elastic modal frequency and exhibits a modulation in a much longer time scale. This slow time modulation over the purely elastic oscillation is due to both the small aerodynamic effects and the small nonlinear friction forces. The correct computation of these two small effects is crucial to determine the final amplitude of the flutter vibration, which basically results from its balance. In this work, we apply asymptotic techniques to consistently derive, from a bladed-disk model, a reduced order model that gives only the time evolution on the slow modulation, filtering out the fast elastic oscillation. This reduced model is numerically integrated with very low computational cost, and we quantitatively compare its results with those from the bladed-disk model. The analysis of the friction saturation of the flutter instability also allows us to conclude that: (i) the final states are always nonlinearly saturated TW; (ii) depending on the initial conditions, there are several different nonlinear TWs that can end up being a final state; and (iii) the possible final TWs are only the more flutter prone ones.

The contrarotating open rotor is, once again, being considered as an alternative to the advanced turbofan to address the growing pressure to cut aviation fuel consumption and carbon dioxide emissions. One of the key challenges is meeting community noise targets at takeoff. Previous open rotor designs are subject to poor efficiency at takeoff due to the presence of large regions of separated flow on the blades as a result of the high incidence needed to achieve the required thrust. This is a consequence of the fixed rotor rotational speed constraint typical of variable pitch propellers. Within the study described in this paper, an improved operation is proposed to improve performance and reduce rotor-rotor interaction noise at takeoff. Three-dimensional computational fluid dynamics (CFD) calculations have been performed on an open rotor rig at a range of takeoff operating conditions. These have been complemented by analytical tone noise predictions to quantify the noise benefits of the approach. The results presented show that for a given thrust, a combination of reduced rotor pitch and increased rotor rotational speed can be used to reduce the incidence onto the front rotor blades. This is shown to eliminate regions of flow separation, reduce the front rotor tip loss and reduce the downstream stream tube contraction. The wakes from the front rotor are also made wider with lower velocity defect, which is found to lead to reduced interaction tone noise. Unfortunately, the necessary increase in blade speed leads to higher relative Mach numbers, which can increase rotor alone noise. In summary, the combined CFD and aeroacoustic analysis in this paper shows how careful operation of an open rotor at takeoff, with moderate levels of repitch and speed increase, can lead to improved front rotor efficiency as well as appreciably lower overall noise across all directivities.

Dynamically scaled experiments and numerical analyses are performed to study the effects of the wake from an upstream wind turbine on the aerodynamics and performance of a downstream wind turbine. The experiments are carried out in the dynamically scaled wind turbine test facility at ETH Zurich. A five-hole steady-state probe is used to characterize the cross-sectional distribution of velocity at different locations downstream of the wake-generating turbine. The performance of the downstream wind turbine is measured with an in-line torquemeter. The velocity field in the wind turbine wake is found to differ significantly from the velocity field assumed in numerical wake models. The velocity at hub height does not increase monotonically up to the freestream velocity with downstream distance in the wake. Furthermore, the flowfield is found to vary significantly radially and azimuthally. The application of wake models that assume a constant axial velocity profile in the wake based on the measured hub-height velocity can lead to errors in annual energy production predictions of the order of 5% for typical wind farms. The application of wake models that assume an axisymmetric Gaussian velocity profile could lead to prediction errors of the order of 20%. Thus modeling wind turbine wakes more accurately, in particular by accounting for radial variations correctly, could increase the accuracy of annual energy production predictions by 5%–20%.

This two-part paper presents a detailed experimental investigation of the laminar separation and transition phenomena on the suction surface of a high-lift low-pressure turbine airfoil, PakB. The first part describes the influence of Reynolds number, freestream turbulence intensity and turbulence length scale on the PakB airfoil under steady inflow conditions. The present measurements are distinctive in that a closely-spaced array of hot-film sensors has allowed a very detailed examination of the suction surface boundary layer behavior. In addition, this paper presents a technique for interpreting the transition process in steady, and periodically unsteady, separated flows based on dynamic and statistical properties of the hot-film measurements. Measurements were made in a low-speed linear cascade facility at Reynolds numbers between 25,000 and 150,000 at three freestream turbulence intensity levels of 0.4%, 2%, and 4%. Two separate grids were used to generate turbulence intensity of 4% with integral length scales of about 10% and 40% of the airfoil axial chord length. While the higher levels of turbulence intensity promoted earlier transition and a shorter separation bubble, turbulence length scale did not have a noticeable effect on the transition process. The size of the suction side separation bubble increased with decreasing Reynolds number, and under low freestream turbulence levels the bubble failed to reattach at low Reynolds numbers. As expected, the losses increased with the length of the separation bubble, and increased significantly when the bubble failed to reattach.

The exit flow field of the fan root of large turbofan engines defines the inlet conditions to the core compressor. This in turn could have significant impact to the performance of the core compressor. This study is aimed to resolve two related issues concerning the impact of the fan root flow on the core compressor performance: to establish the effect of an increased loss at the inlet on the engine specific fuel consumption (SFC) and to assess the effect of the radial distribution of the fan root flow on the engine performance. With understanding of these issues, the geometric parameters and design details which can produce a more uniform core flow at the exit of the fan stage module can be identified. The fan root flow is analyzed with methods of different complexity and fidelity. A simple cycle analysis is used to assess the impact on engine SFC of a stagnation pressure deficit at the fan root; a throughflow code is used for the preliminary study of the curvature effect of the root flow path, and 3D RANS CFD calculations are then used to simulate the flow path from the inlet of the fan to the first stage of the core compressor. The adequacy of the application of the numerical code in this case has been assessed and confirmed by the comparison with the experimental data for two rig configurations. The results of this study show that the flow at the fan hub region is very complex and dominated by 3D effects. The interaction of the secondary flow with real geometries, such as leakage flows, is found to have a strong detrimental effect on the core performance. The curvature of the hub end wall is a key parameter controlling the fan root flow topology; it influences the strength of the secondary flow, the spanwise distribution of the flow, and its sensitivity to leakage flow. With this understanding, it is possible to redesign of the fan hub flow path to reduce the loss generation by a significant amount.

A frequency-domain method has been developed to predict and comprehensively analyze the limit-cycle flutter-induced vibrations in bladed disks and other structures with nonlinear contact interfaces. The method allows, for the first time, direct calculation of the limit-cycle amplitudes and frequencies as functions of contact interface parameters and aerodynamic characteristics using realistic large-scale finite element models of structures. The effects of the parameters of nonlinear contact interfaces on limit-cycle amplitudes and frequencies have been explored for major types of nonlinearities occurring in gas-turbine structures. New mechanisms of limiting the flutter-induced vibrations have been revealed and explained.

The effect of the structural coupling in the aeroelastic stability of a packet of low-pressure turbine vanes is studied in detail. The dynamics of a 3D sector vane is reduced to that of a simplified mass-spring model to enhance the understanding of its dynamics and to perform sensitivity studies. It is concluded that the dynamics of the simplified model retains the basic features of the finite element three-dimensional model. A linear fully coupled analysis in the frequency domain of the 3D vane sector has been conducted. It is concluded that the small structural coupling provided by the casing and the inter-stage seal is essential to explain the experimental evidences. It is shown that the use of fully coupled aero/structural methods is necessary to retain the mode interaction that takes place in this type of configurations.