Improving fan design and performance prediction requires an understanding of the impacts of inlet distortion. Using full-annulus unsteady Reynolds-averaged Navier–Stokes (URANS) simulations, the effects of hub and tip radial distortions on the Air Force Research Laboratory (AFRL) Rotor 4 fan are analyzed. Overall performance, total pressure (Pt) distortion transfer, and total temperature (Tt) distortion generation are presented and compared for three operating points: near-stall, design, and choke. Results from twelve simulations are presented and analyzed. Hub radial inlet distortion magnitudes of 15% and 20%, and a tip radial distortion of 15% magnitude were investigated. The 15% hub radial profile increased stall margin by 18.45% compared with the clean inlet. The 15% tip radial profile decreased the stall margin by 89.47%. An intensity distortion descriptor is used to quantify levels of distortion transfer and generation through the fan. It is shown that hub radial profiles resulted in more significant Pt distortion attenuation and less Tt distortion generation through the rotor than the tip radial profile.

Meanline performance models have a long history of use in the design of radialflow pumps and compressors. Most data used to create correlations are point data such as static pressures and temperature measurements at walls with occasional traverses for more details of the flow field. These point measurements along with limited traverses, while very useful, do not provide details of specific flow characteristics internal to the machine. The Two Zone model, developed by Japikse and others, is a meanline model based on point measurements that enables prediction of machine performance. The model assumes that the flow at the impeller exit is composed of a primary zone that is isentropic and a secondary zone where losses exist. The Two-Zone model can accurately predict power when the fraction of mass flow in the secondary zone (chi) and the deviation angle in the primary (delta(2p)) and secondary (delta(2s)) zones are correctly inferred from the available data. However, it is difficult to ensure that the parameters inferredfrom the Two-Zone flow model structure are uniquely correct. In this paper, the specific details of the impeller exit flow fields required to generate an accurate prediction using the Two-Zone model are extracted from CFD simulations. Using CFD, flow fields at the impeller exit are obtained and analyzed to calculate the model parameters for 10 radial flow machines. The derived two zone parameters are then used to predict the power for each machine. The Two-Zone predictions match the experimental data with an average error of 3.5%.

Aircraft engine performance and stability is limited by the onset of flow instabilities, such as rotating stall or surge. Although much research has been conducted on the subject, the mechanisms by which spike-type stall inception occurs in highspeed machines are still not well understood. This paper investigates the mechanisms of stall inception in a highly-loaded transonic fan rotor using full-annulus Unsteady Reynold’s-Averaged Navier-Stokes (URANS) simulations. Comparisons are made between the flow phenomena at the near-stall operating point and as stall inception occurs. The mechanisms of stall inception are shown to be interactions between the detached bow shock and the tip leakage vortex. These interactions result in two vortices within the blade passage near the rotor tips. The location and strength of these vortices affect leading edge spillage in the adjacent blade row. At the near-stall operating point, the vortices impact the pressure surface of the adjacent blade and then convect downstream. Stall inception occurs when the bow shock has moved far enough upstream to allow the resultant passage vortices from the shock/tip leakage vortex interaction to pass in front of the leading edge of the adjacent blade. After this, radial vortical structures are seen at the blade tips, much like what is observed in low-speed compressors. Understanding the mechanisms of stall inception in high-speed fans will allow for appropriate steps to be taken to delay stall inception and extend the stable operating range of the fan.

A stability control device (SCD) is a passive inlet cover treatment that can be applied to high suction performance inducers. Significant improvements in stability have been observed when an inducer operates with an SCD, including suppression of backflow at the leading edge of the inducer at low off-design flow coefficients. This is possible because of a local increase in mass flow at the leading edge of the pump, which allows an inducer to operate with an incidence near the design point value over a wide range of flow coefficients. In this paper, the suction performance of several inducers was explored with several different SCD geometries and at different flow coefficients. Specifically, five different SCD geometries were considered to explore the influence that SCD bleed slot width and resistance to flow through the SCD channel have on inducer performance. Further, removal of tangential velocity of the energized fluid transporting through the SCD channel was considered for some scenarios to highlight the impact of swirl on inducer performance. The results reveal that for all inducers and SCD combinations considered, the most important factor affecting the mass flow through the SCD, and subsequent mass flow gain, is the inlet diffusion of the inducer. This holds for both single-phase and multiphase scenarios. Further, the cavitation number where cavitation first starts to develop in the blade throat passage of the inducer is primarily dependent on the inlet blade angle and not the SCD geometry. Consequently, the shape of the cavitation breakdown curve is largely determined by the inducer blade angle.

The effect of distorted inflow on jet engine fan stability is an important consideration in both commercial and high performance aircraft applications. Novel commercial aircraft configurations offer significant reductions in fuel burn, but require engine fans to continuously operate with severe amounts of inlet distortion. Complex inlet ducts in high performance aircraft also result in distorted flow at the fan inlet. This paper investigates and compares the mechanisms of stall inception for cases of both clean and distorted inlet flow. Unsteady RANS simulations of stall inception in a high performance transonic fan were obtained for both cases. In both instances, the mechanism for stall inception is shown to be interactions between the detached bow shock and the tip clearance vortex, which causes the formation of two vortices within the blade passage. The location and strength of these vortices both affect the LE spillage into the adjacent blade rows. When inlet distortion is present, mass redistribution upstream of the fan results in variations in rotor incidence. An area of high incidence is located outside the circumferential extent of the applied distorted sector. In the high incidence region, the shock is detached 3.9%–8.1% chord more than the clean inlet case, making LE spillage more likely. The same mechanisms of stall inception are observed when distortion is present. However, the severity of the stalling phenomena is more pronounced within the high incidence region of the annulus. A rotating stall cell grows out of the stalled passages present at the near-stall operating point and ultimately extends 59% chord upstream of the LE and 18.7% span radially. Understanding the effect of distortion on the mechanisms of stall inception will allow appropriate steps to be taken to extend the stable operating range of modern commercial and high performance fans.

An increased understanding of how inlet flow distortion affects transonic fans enables improved fan design and performance prediction. Unsteady Reynolds-Averaged Navier Stokes (URANS) simulations were carried out to understand how the transition abruptness between the clean and distorted sectors in the inlet flow as well as the circumferential extent of the distorted sector affects the distortion transfer and generation throughout the fan. Simulations on two main distortion sector sizes were carried out. For each sector size, variants of decreasing levels of transition abruptness were applied to the inlet of the Rotor 4 geometry. Simulations were conducted at various operating points, ranging from choke to near-stall. Fourier-based distortion descriptors were used to quantify levels of distortion transfer and generation at various axial locations. A less abrupt transition reduces the amount of distortion transfer by 23.4% at near-stall, the most of any operating point. At all operating points, both the incidence and the shock strengths are diminished by a less abrupt transition, which causes a reduction in the variation of local rotor power. This is most pronounced at near-stall, where the maximum rotor power is decreased by 11.8% (1.25 kW) and 4.6% (1.24 kW) respectively at 10% and 90% span. The size of the distorted sector affects the amount of distortion transfer and generation. This is especially true near the rotor tips at near-stall, where a high total pressure sector extends 33% less circumferentially for 135° shapes than for 90° shapes. This is shown to be due to the dynamic response of the fan.

Fatigue life analysis and FEA require high computational costs for aerospace models. Goodman diagrams are useful tools for determining the fatigue life of a part, and depend on structural and modal FEA simulations. Surrogate models have been used in fatigue analysis to predict the steady stresses and alternating stresses FEA model for every par node enabling a dynamically-updated Goodman diagram that responds to design parameter changes in real time. This research analyzes a jet engine compressor blade, and compares the accuracy of these nodal surrogate methods to previous surrogate methods where only a single value is predicted for failure analysis. Quantitative data summarizing the error between the single value and full field models are presented for the transonic Purdue blade.

Computational cost of high-fidelity simulations limits the number of evaluations which may be performed in design exploration and optimization. Surrogates based on samples of multiple fidelities are used to decrease computational cost and lower error from single-fidelity surrogates. This paper develops a novel multi-fidelity surrogate model based on principal components which are shared between multiple fidelities of finite element model samples. This method does not require a common grid between the fidelities, further reducing computational cost. The new method was tested on various design spaces of the Transonic Purdue Research Compressor and compared to other common and novel multi-fidelity methods. The new method was more accurate and required less computational cost than the other tested methods. Little to no increase in computational cost was needed to reduce surrogate error to 50% of the single-fidelity error. For fixed error, the computational cost was reduced by more than 75%. These results were also validated by testing the method on a more complex turbomachinery blade, Parametric Blade Study Rotor 4. The decreased error and computational cost improve effectiveness of design exploration and optimization. Such improvements help meet the demand for cleaner and safer engines by allowing high-fidelity design exploration within reasonable time frames.

There is widespread interest in using pressure gain combustion in gas turbine engines to increase gas turbine engine efficiency and reduce fuel consumption. However, the fluctuating turbine inlet conditions inherent with pressure gain combustion cause a decrease in turbine efficiency. Designing a turbine for pulsing flow would counteract these losses. An optimization of turbine geometry for pulsing flow was conducted with entropy generation as the objective function. A surrogate model was used for the optimizations based on data extracted from two-dimensional computational fluid dynamics simulations. Optimizations run for different pulsing amplitudes informed a revised turbine design. The new turbine geometry was validated with a periodic, time-accurate simulation, and a decrease in entropy generation of 35% was demonstrated. The design recommendations were to weight the design of the turbine toward the peak of the pressure pulse, to consider the range of inlet angles and decrease the camber near the leading edge, and to reduce the blade turning.

Optimization of the structural response for a part relies upon computationally expensive simulations such as finite element analysis (FEA). Surrogate models are able to make cheap predictions of the results; however, traditionally they can only predict a single value (SV) such as a maximum stress or weight value. Recently, a surrogate modeling method has been developed for predicting the full field (FF) of nodal responses in an FEA simulation. This research applies FF surrogate models to optimization, and explores various techniques that are uniquely enabled by these cheap-yet detailed-predictions. Because the FF surrogate models predict the response at every node, constraints and objectives can be spatially defined on a part rather than act on a single value. Regional constraints allow for control over a subset of the nodes in areas of interest on the part. Location-based objectives find designs that either draw a response closer to or drive a response away from a particular node. Pattern-matching objectives find designs in the design space that have a response pattern across the part surface that is as similar as possible to a pre-defined response pattern. These techniques extend the usefulness of FF surrogate models as well as optimization of FEA results for exploring a design space and improving a design.

Design space exploration (DSE) is a systematic analysis of designs based on parameters that are varied. DSE, however, often involves computationally expensive analyses that require large amounts of time and computer resources. To overcome this obstacle, tools and methods have been developed that allow the designer to quickly explore the design space. One such tool is surrogate modeling. A surrogate model's predictions are based off the relationship between the inputs and the outputs. If this relationship is sufficiently nonlinear, a nonlinear dimensionality reduction technique may unfold the data and create a linear relationship by discovering the data's true dimensionality. The resultant surrogate has the potential to be more accurate due to the simplified relationship between the inputs and outputs. This paper uses surrogate models based on dimensionality reduction to predict finite element analysis (FEA) nodal properties of a jet engine compressor blade. Four different dimensionality reduction methods are compared, namely PCA, KPCA, ISOMAP, and LLE. The results show that nonlinear dimensionality-reduction-based surrogate models can reduce surrogate error in sufficiently nonlinear data spaces. To measure the linearity of the manifold, a new term, called the manifold distance ratio (MDR), is introduced. Nonlinear dimensionality-reduction-based surrogate models can have less error than PCA-based surrogate models when the MDR >1.7. For our application, ISOMAP-based surrogate models decreased the mean error of the surrogate for predicting nodal stresses by 35.7% compared to PCA. The lowest error in predicting the nodal coordinates was the inverse-transform-based PCA. While all errors in nodal coordinates were small, less than 0.23%, the other linear and nonlinear algorithms had between 2.4 and 3.2 times more error.

Design space exploration (DSE) is the process whereby a designer seeks to understand some results across a set of design variations. Structural DSE of turbomachinery compressor blades is often challenging because the large number of design variables make it difficult to learn the effect that each variable has upon the stress contours. Principal component analysis (PCA) of the stress contours is used as a way to understand how the stress contours change over the design space. Two methods are introduced to address the challenge of understanding how the stress changes over a large number of variables. First, a two-point correlation is applied to relate the design variables to the scores of each principal component. Second, a coupling of the stress and coordinate location of each node in PCA is developed which also indicates how the stress variations relate to geometric variations. These provide insight to how design variables influence the stress. It is shown how these methods use PCA as DSE tools to better explore the structural design space of compressor blades. Better DSE can improve compressor blades and the computational cost needed for their design.

One of the challenges of integrating pressure gain combustion into a gas turbine engine is that a turbine driven by pulsing flow experiences a decrease in efficiency. Computational fluid dynamic simulations validated with experiments showed that pulse amplitude is the driving factor for decreased turbine efficiency and not the pulsing frequency. A quadratic correlation between turbine efficiency and corrected pulse amplitude is presented. Incidence variation is shown to cause the change in turbine efficiency and a correlation between corrected incidence and corrected amplitude is shown to predict turbine efficiency.

Full annulus URANS simulations are presented to investigate distortion phase shift in a single stage transonic fan. The fan is subject to a 90 degree sector inlet total pressure distortion. Simulation results are presented for choke, design, and near-stall operating conditions. Circumferential profiles of swirl, total pressure, total temperature, power, and phase shift are analyzed at 10%, 30%, 50%, 70%, and 90% span. Phase shift is a measure of the rotational translation of a distortion profile and is valuable for understanding the translational motion of distortion as it passes through the fan. At choke and design operating conditions the phase shift is positive in the direction of rotor rotation. At near-stall the phase shift is negative opposite rotor rotation for the outer 30% span. Local power reaches a maximum value, resulting in the generation of a uniform pressure region. This region causes increased distortion content in the near-stall case over a wide circumferential extent resulting in a negative phase shift.

Characterization of distortion transfer and generation through fans with distorted inlet conditions enables progress towards designs with improved distortion tolerance. The abruptness of transition from undistorted to distorted total pressure regions at the inlet impacts the induced swirl profile and therefore the distortion transfer and generation. These impacts are characterized using URANS simulations of PBS Rotor 4 geometry under a variety of inlet distortion profiles. A 90° and a 135° sector, both of 15% total pressure distortion, are considered. Variants of each sector size, with decreasing levels of distortion transition abruptness, are each applied to the fan. Fourier-based distortion descriptors are used to quantify levels of distortion transfer and generation at axial locations through the fan, principally at the stator inlet. It is shown that a gradual transition in distortion at the inlet results in decreased levels of distortion transfer and generation. The flow physics resulting in this reduction are explored.

Turbomachines are an integral part of society, with global trends demanding more efficient designs while staying within structural limits. Fan blade designs must be designed for both steady and vibratory structural responses. Design space exploration (DSE) of turbomachinery blades allows improved designs to be found. DSE requires samples of vibratory responses. Traditionally, analysis to obtain these samples is too computationally expensive for thorough DSE. This work develops a simplified analysis method based on Reynolds-averaged Navier-Stokes (RANS) computational fluid dynamics (CFD) and harmonic mode superposition (HMS) finite element analysis (FEA). This reduces the computational cost and allows for enough samples to create surrogate models. This work also develops a surrogate method which indirectly emulates the vibratory responses to accurately handle the large spikes in vibratory stress found throughout the design space. It was found that combining these methods allows for accurate emulation of a fan blade design space while accounting for vibratory stress. The surrogates with improved accuracy allow better designs to be found while ensuring that those designs meet structural requirements.

An inducer is used as the first stage of high suction performance pump. It pressurizes the fluid to delay the onset of cavitation, which can adversely affect performance in a centrifugal pump. In this paper, the performance of a water pump inducer has been explored with and without the implementation of a stability control device (SCD). This device is an inlet cover bleed system that removes high-energy fluid near the blade leading edge and reinjects it back upstream. The research was conducted by running multiphase, time-accurate computational fluid dynamic (CFD) simulations at the design flow coefficient and at low, off-design flow coefficients. The suction performance and stability for the same inducer with and without the implementation of the SCD has been explored. An improvement in stability and suction performance was observed when the SCD was implemented. Without the SCD, the inducer developed backflow at the blade tip, which led to rotating cavitation and larger rotordynamic forces. With the SCD, no significant cavitation instabilities developed, and the rotordynamic forces remained small. The lack of cavitation instabilities also allowed the inducer to operate at lower inlet pressures, increasing the suction performance of the inducer.

Understanding distortion transfer and generation through fan and compressor blade rows is able to assist in blade design and performance prediction. Using full annulus URANS simulations, the effects of distortion as it passes through the rotor of a transonic fan at five radial locations (10%, 30%, 50%, 70%, and 90% span) are analyzed. The inlet distortion profile is a 90-degree sector with a 15% total pressure deficit. Fourier distortion descriptors are used in this study to quantitatively describe distortion transfer and generation. Results are presented and compared for three operating points (near-stall, design, and choke). These results are used to explain the relationship between inlet total pressure distortion, pressure-induced swirl, total pressure distortion transfer, total temperature distortion generation, and circumferential rotor power variation. It is shown that very large changes in pressure-induced swirl and distortion transfer and generation occur between near-stall and design, but only small changes are seen between design and choke. The greatest changes are shown to be near the tip. Local power variations are shown to correlate with total pressure distortion transfer and total temperature distortion generation.

Inducers are used as a first stage in pumps to minimize cavitation and allow the pump to operate at lower inlet head conditions. Inlet flow recirculation or backflow in the inducer occurs at low flow conditions and can lead to instabilities and cavitation-induced head breakdown. Backflow of an inducer with a tip clearance (TC) of τ = 0.32% and with no tip clearance (NTC) is examined with a series of computational fluid dynamics simulations. Removing the TC eliminates tip leakage flow; however, backflow is still observed. In fact, the NTC case showed a 37% increase in the length of the upstream backflow penetration. Tip leakage flow does instigate a smaller secondary leading edge tip vortex that is separate from the much larger backflow structure. A comprehensive analysis of these simulations suggests that blade inlet diffusion, not tip leakage flow, is the fundamental mechanism leading to the formation of backflow.