Full Annulus Research Articles

A Computational Study of Temperature Driven Low Engine Order Forced Response In High Pressure Turbines

Abstract This paper reports the results of computational studies of the effect of combustor exit temperature distortions on Low Engine Order (LEO) forced response of a High Pressure Turbine (HPT). Forced response of this kind occurs at frequencies below the Stator Vane Passing Frequency (SVPF) and can be a major cause of High Cycle Fatigue (HCF) in turbines due to its tendency to excite fundamental modes of vibration. This paper investigates the extent through which temperature distortions act as a forcing stimulus in HPT rotor rows, through measuring unsteady pressure and modal force magnitude recorded from full annulus unsteady simulations of the MT1 stage: a low temperature, unshrouded, HPT rig. Rotor relative incidence angle variations are shown to be the key mechanism through which temperature acts as a forcer in HPT rotor rows, whilst temperature driven forced response is shown to be dependent on the magnitude of the modal content of the upstream temperature waves. These findings are used to build a reduced domain tool for blocked burner forced response prediction, which is shown to be accurate to an RMS error of 2.66%, far beyond the current accepted standard for forcing prediction of this kind.

Unsteady numerical study on the effects of inlet distortion parameters on the aerodynamic instability of the centrifugal compressor

Total pressure inlet distortion significantly impairs the aerodynamic performance and flow stability of centrifugal compressors. This paper presents a full annulus unsteady numerical simulation to investigate the flow mechanisms under varying inlet distortion parameters. First, the flow characteristics of a uniform inflow is studied as a baseline for comparison with distorted inlet conditions. Results indicate that the interaction between shock waves and tip leakage flow is a primary factor leading to aerodynamic instability. The instability signal in this region is detected through fast Fourier transform and frequency slice wavelet transform (FSWT). Under near stall condition, different distortion parameters are set to research the effects on aerodynamic performance, flow mechanisms, and instability characteristics. The results reveal that distortion intensity has the most significant impact, causing a maximum stall margin loss of 55.43%, followed by distortion angle with a maximum stall margin loss of 43.02%. Total pressure inlet distortion leads to a deterioration in aerodynamic performance, primarily due to premature occurrences of unstable flow phenomena such as leading-edge spillage and trailing-edge backflow. The onset key features triggering aerodynamic instability are identified as leading-edge spillage vortex, tip leakage vortex, and passage vortex. The continuous disintegration of the tip leakage vortex results in low-frequency fluctuating energy exhibiting multipeak characteristics, with pulsation peaks centered around 0.5 BPF, related to the spike stall of the compressor. The high-energy frequency band dissipates over time in the time–frequency spectrum, as shown by FSWT results, indicating the characteristics of instability in the flow.

Inductive plasma excitation forcing configuration on reduction of tip distorted inflow effect on the aerodynamic stability of axial compressor rotor

This research delves into the mitigating impacts of dielectric barrier discharge (DBD) plasma excitation induced forcing orientation against the detrimental consequences of distinct radial tip distortions which in turn affect the axial compressor rotor performance and alters the flow structure at the tip region. Full annulus transient CFD simulation was utilized to evaluate the consequences of plasma actuation at distorted conditions with different blockage percentages. Beyond flow field and frequency analysis, the study further characterized rotor performance under different conditions by evaluating key performance metrics, including total pressure rise coefficient, stall margin variation, and span-wise rotor inlet velocity distribution. The injection of momentum caused by plasma actuators to the low-energy region behind the distortion screens proved to be effective on rotor aerodynamic stability facing radial tip distortion. In the case where 15 % of the inlet area was blocked, the stall margin varied from -8 % to -3.5 % with axial plasma actuators in action. However, the best configuration of plasma actuators for the enhancement of the stall margin and flow characteristics was identified to have opposite forcing direction with respect to the rotor rotational velocity. Additionally, these actuators suppressed frequencies caused by fluctuations in the rotor blade row tip leakage vortex, suggesting an improvement in the flow pattern within the rotor tip area.

Assessment of a body force modeling approach to examine the aerodynamics of high bypass ratio fan stages

The primary focus of modern civil jet engine intake design is to enhance propulsive efficiency through an increase in the bypass ratio of the turbofan engine, which requires an enlargement in fan diameter. However, this enlargement results in increased weight and drag due to the need for a larger nacelle to surround the fan. To mitigate these issues, efforts are being made to shorten the axial length of the aircraft engine intakes. Nonetheless, this reduction leads to more vulnerable fan aerodynamics since the diffusion within the intake must occur over a shorter distance, causing a higher possibility of flow separation. Common intake design methods employ throughflow nacelle models that do not consider the aerodynamic effects of the fan, including blockage, mass flow redistribution, and suction. The utilization of a full annulus domain numerical setup, which is necessary to capture inlet distortions, demands excessive computational resources particularly during the design phase. As a result, reduced order models, such as the body force model, can be utilized to simulate fan intake aerodynamics. This paper will assess the body force model and its abilities by comparing conventional bladed single passage simulations with the body force approach.

Aerodynamic and aeroacoustic design of electric ducted fans

This paper describes the aerodynamic and aeroacoustic design of an electric ducted fan (EDF). A prototype EDF was designed and built with three different rotor-stator sets to operate at design flow coefficients ϕd=[0.60,0.75,0.90]. Computational Fluid Dynamics (CFD) simulations show that changing the design flow coefficient changes the proportions of endwall and blade profile loss, with the ϕd=0.75 design providing the optimum balance. CFD predictions of aerodynamic performance are compared with experimental measurements and show good agreement. Data collected from full annulus Unsteady Reynolds-Averaged Navier Stokes (URANS) CFD simulations is used to predict tonal noise radiation using a numerical solution to Farassat's Formulation 1A, and compare favourably to acoustic measurements in an anechoic chamber. All three EDF iterations show similar levels of broadband noise radiation, although tonal noise radiation decreases with increasing design flow coefficient and results in more favourable Sound Quality Metrics performance. Design flow coefficient is shown to have a first-order impact on both the aerodynamic and aeroacoustic behaviour of the EDF, and it is proposed that an optimal design, encompassing both aerodynamic performance and reduced acoustic impact, can be achieved at design flow coefficients between ϕd=0.75 and ϕd=0.90.

Numerical Investigation of the Excitation Characteristics of Contaminated Nozzle Rings

The deposition of combustion residues in the nozzle ring (NR) of a turbocharger turbine stage changes the NR geometry significantly in a random manner. The resultant complex and highly asymmetric geometry induces low engine order (LEO) excitation, which may lead to resonance excitation of rotor blades and high cycle fatigue (HCF) failure. Therefore, a suitable prediction workflow is of great importance for the design and validation phases. The prediction of LEO excitation is, however, computationally expensive as high-fidelity, full annulus CFD models are required. Previous investigations showed that a steady-state computational model consisting of the volute, the NR, and a radial extension is suitable to reduce the computational costs massively and to qualitatively predict the level of LEO forced response. In the current paper, the aerodynamic excitation of 69 real contaminated NRs is analyzed using this simplified approach. The results obtained by the simplified simulation model are used to select 13 contaminated NR geometries, which are then simulated with a model of the entire turbine stage, including the rotor, in a transient time-marching manner to provide high-fidelity simulation results for the verification of the simplified approach. Furthermore, two contamination patterns are analyzed in a more detailed manner regarding their aerodynamic excitation. It is found that the simplified model can be used to identify and classify contamination patterns that lead to high blade vibration amplitudes. In cases where transient effects occurring in the rotor alter the harmonic pressure field significantly, the ability of the simplified approach to predict the LEO excitation is not sufficient.

Resolving nonuniform flow in gas turbines: challenges, progress, and moving forward

Abstract The flow in gas turbines exhibits a highly unsteady, complex and nonuniform manner, which presents two main challenges. Firstly, it introduces instrumentation errors, contributing to uncertainties when calculating one-dimensional performance metrics during rig or engine tests using fixed-location rakes. Secondly, it raises mechanical concerns, including high-cycle fatigue due to blade row interactions in turbomachines and thermal fatigue caused by hot-streaks at the combustor exit. Experimental characterisation of the flow nonuniformity in gas turbines is highly challenging due to the confined space and harsh environment for instrumentation. This paper presents recent efforts to address this issue by resolving the nonuniform flow in gas turbines using spatially under-sampled measurements. The proposed approach utilises discrete probe data and leverages a ‘Fourier-based approximation’ method developed by the author. The technique has undergone preliminary experimental validation involving reconstruction of the total pressure distribution in a multi-stage axial compressor and the total temperature field at the exit of the combustor and high-pressure turbine. Results show that, in the multi-stage axial compressor environment, reconstruction of inter-stage total pressure is achieved using a reduced dataset covering less than 20% of the annulus with reasonably good accuracy. The reconstructed total pressure yields almost identical mean total pressure values at all spanwise locations, with a maximum deviation of less than 0.02%. Additionally, reconstruction of the total temperature distribution at an engine-representative full annulus combustor is achieved using measurements at ten carefully selected circumferential locations. Results show that the reconstructed temperature field successfully captures the primary features associated with combustor exit temperature flow. The reconstructed temperature field yields excellent agreement in the magnitude of radial temperature distribution factor (RTDF) and overall temperature distribution factor (OTDF) predictions to the experiment, with a deviation of less than 0.5% for RTDF and less than 2.5% for OTDF. Lastly, reconstruction of the total temperature distribution at the exit of the GE E3 high-pressure turbine (HPT) is achieved using measurements at eight carefully selected circumferential locations. Results demonstrate remarkable robustness in resolving the temperature profile at the HPT exit with high fidelity, irrespective of the HPT inlet conditions. The initial validation results are promising, demonstrating that the new probe layout scheme and the Fourier-based approximation method enable effective characterisation of flow nonuniformity in gas turbines, thereby providing valuable insights into the complex flow of gas turbine engines.

Strategies for High-Accuracy Measurements of Stage Efficiency for a Cooled Turbine

Abstract Gas turbines are used in a broad range of aerospace and land-based applications from power generation to aviation, and their usage is projected to continue to grow. It is therefore critical to improve turbine efficiency thereby reducing fuel consumption and carbon emissions. Demonstration of new efficiency-increasing technologies requires efficiency measurements that are accurate and repeatable. The Steady Thermal Aero Research Turbine (START) Laboratory at the Pennsylvania State University uses a unique 360-deg traversing system for temperature and pressure probes with redundant torque measurements to quantify thermal efficiency for a single-stage cooled test turbine. Flows in the full annulus have been analyzed and compared with subsector traverse segments centered at different circumferential positions to determine the appropriate sector size. The results from this investigation indicate that the full 360-deg measurement is recommended to minimize variation in calculated stage efficiencies. This study also compares the circumferential variations in thermodynamic and mechanical efficiency definitions, finding that the thermodynamic efficiency calculation results in a higher accuracy for full exit plane measurements. A statistical analysis was then performed to determine the number of 360-deg traverse measurements required to achieve a precision uncertainty at most that of the bias uncertainty. This study establishes guidelines to streamline experimental procedures by limiting the necessary test count per day per operating condition to five measurements for at most four test days. Following these procedures establishes a bias of ɛb = 0.19 points, resulting in a total uncertainty of at most ɛt = 0.32 points.

Simulation and validation of instability transient process in multistage high-speed axial compressor based on the body-force model

In this paper, a method for simulating the instability transient process of the axial compression system based on the body-force model is developed, and a corresponding simulation program is developed. Simulations of the transient process of instability were carried out on a high-speed four-stage compressor and compared with experimental data. At 50% of the design rotational speed, the type of instability was rotating stall, and the simulated and experimental stall cell propagation speed were very close to each other. At 70% of the design rotational speed, the type of instability was surge. A “surge loop” was simulated, and the surge period and the percentage of time spent in each phase were consistent with the experiments. The simulation successfully predicted the blockage in the surge re-pressurization phase, proving the reliability of the simulation results. In addition, the computation yields more information about the flow field. By summing the blade forces of all grids on a blade row by volume, the surge loadings are obtained. The analysis of the axial momentum equation shows that the obtained blade force variations are reasonable. The simulation time of the multistage axial compressor is greatly reduced compared to the full annulus three-dimensional unsteady Reynolds-averaged Navier–Stokes method, demonstrating its great advantage in the design phase of the compressor.

The impact of second-harmonic tip clearance distributions caused by casing Ovalization and blade height variation in a large-scale compressor stage

Affected by extreme operating conditions, external loading and casing stiffness, the casing ovalization will potentially occur for the axial compressor, resulting in a second harmonic distribution of rotor tip clearances over the full annulus. The same clearance layout can be reproduced by varying the blade height over the circumference with a circular casing. In this paper, both of these non-axisymmetric configurations are respectively modelled with full-annulus simulations in the first 1.5 stage of the large-scale research compressor at Shanghai Jiao Tong University (SJTU-LSRC), with the blade adjustment item further complemented by experimental measurements.The impacts are evaluated and compared in terms of compressor performance and flow mechanisms. Ovacasing case leads to a 32.57 % stall margin (SM) benefit compared to the prototype, while ovablade experiences a 31.75 % SM extension. Both configurations cause a 4 % efficiency decrease at the design conditions. The second-harmonic clearance distribution results in a 15 lag in ovablade and a 30 lag in ovacasing for the total pressure ratio at the rotor exit, due to local mass flow adaptation related to circumferential tip clearances. Ovacasing configuration exhibits substantial reverse flow and a low-velocity region in the tip of large clearance sector, while ovablade shows reduced reverse flow and secondary leakage flow. This underscores the impact of blade tip relationships and height differences from adjacent rotors on leakage flow and its secondary behavior, even with identical clearances. There exists a critical clearance (2.57 % span), below which adjusting blade heights effectively simulates the performance impact and flow field redistribution caused by casing ovalization and rotor eccentricity.

Numerical Modeling to Investigate the Aerodynamics of a Boundary-Layer Ingested Transonic Fan

Boundary-layer ingested engines have the potential to offer significantly reduced fuel burn, but the fan stage must be designed to run efficiently with a distorted inflow. It must also be able to withstand unsteady aerodynamic loads resulting from a complex nonuniform flowfield. This paper applies different numerical methods for an improved understanding of the aerodynamic interaction between a transonic fan and inlet distortion. A single-stage transonic tail cone thruster fan was designed using both in-house and commercial tools operating in an inlet distortion flowfield. This paper demonstrates that the relevant metrics required to compute the aerodynamic performance of a fan stage in distorted conditions can be reasonably modeled with a few harmonics using the nonlinear harmonic method in a fraction of time in comparison to a full annulus time marching solution. The nonlinear harmonic method also reduces the computational domain, and hence reduces the solution runtime by an order of magnitude. However, it fails to accurately resolve the wake and potential field transfer across the blade rows due to a limited number of harmonics being applied. A detailed aerodynamic description of the unsteady inflow distortion, the interacting blade-row mechanisms, the flow redistribution upstream of the rotor, the distortion transfer across the different blade rows, and the corresponding aerodynamic losses can be analyzed accurately using only a full annulus time-marching method.

Numerical study on the effect of distortion of S-duct on flow field and performance of a full annulus transonic fan

Abstract To reveal the influence of distortion of S-duct on the flow field and performance of transonic fan, full annulus unsteady numerical simulation was carried out under S-duct/fan integrated condition. This study focuses on the coupling flow of S-duct/fan integrated condition under peak efficiency (PE) condition and near stall (NS) condition. Results show that, compared with the condition that S-duct or fan is investigated alone, the distortion of S-duct is suppressed under S-duct/fan integrated condition. The curved structure of S-duct is the important reason for the flow migration from undistorted region to distortion region. Higher inlet relative Mach number and more work input of the rotor are observed in the counter-swirl region. The flow separation in S-duct is weakened under NS condition compared with that of PE condition. Counter-swirl region of blade tip has larger region of blockage, and rotating stall inception is most likely to occur in this region.

Numerical investigations of mass-flux redistribution and performance prediction model on coupling effects of total pressure distortion-low Reynolds number in axial compressor

Stability assessment is one of the important subjects in airworthiness certification. Many destabilizing factors can deteriorate the performance and stability margin of the compressor, and there are coupling effects between the external factors. This paper applies a computational method to investigate the coupling effects of total pressure distortion and low Reynolds number on an axial transonic compressor. The study was carried out on NASA stage 67, in which extensive open experimental data is available. Full annulus, steady, three-dimensional CFD has been used to study the coupling effects of 180-deg inlet distortion at low Reynolds number. The assessment of the stability margin loss under coupling effects by traditional methods results in a relative error of 27%. The local operating point of the rotor at different passages has been tracked, and the results reveal the reasons for the failure of model predictions. Combined with the analysis of the flow field under coupling effects, development solutions for the traditional model are proposed. The modifications to the model agree well with the calculation results. Finally, the improved model was validated. And according to the analysis of the rotor passage working points, the rationality of the model modification was verified.

Using Tip Injection to Stability Enhancement of a Transonic Centrifugal Impeller with Inlet Distortion

This paper aims to understand the effects of circumferential inlet distortion and tip injection on a transonic impeller performance and flow field. For distorted inflow, the impeller is subjected to a stationary 120-degrees circumferential total pressure distortion. Full annulus unsteady three-dimensional analysis has been used to study the inlet distortion and tip injection effects on the impeller performance, stability and flow field. The results show that the circumferential inlet distortion reduces the impeller total pressure ratio and adiabatic efficiency; however, it has no significant impact on the safe operating range. Unlike the inlet distortion, the tip injection considerably increases the operating range. According to the results, the distortion and tip injection effect on the compressor performance is mainly due to changes in tip leakage flow. The inlet distortion has unfavorable influences on the flow field, especially near the impeller tip; however, the tip injection ameliorates the flow field in this region. In both the clean and distorted inflow, the tip injection causes downstream shock transmission, weakening the shock-tip leakage interaction. Hence, stall inception is postponed, and the impeller stability is improved in the presence of the tip injection.

Fan Stability Enhancement With Partial Casing Treatments

Abstract Casing treatments are used to extend the stability margin of fans or compressors. These typically have an axisymmetric arrangement, but for a fan exposed to non-axisymmetric distortion, a casing treatment applied to a fraction of the annulus is of interest to target a particular distortion feature. In this paper, a partial casing treatment has been designed with axial-skewed slots, and an experimental investigation has been carried out to determine the effects of the treatment extent and position. Three inlet conditions were tested for each extent of casing treatment: clean, radial tip-low distortion, and half-annulus tip-low distortion. For axisymmetric inlet conditions, the stall margin and efficiency penalty increased approximately linearly with casing treatment extent. However, for half-annulus tip-low distortion, there was a strong dependence between the position of the casing treatment relative to the distorted sector and the stall margin improvement. If the casing treatment is positioned where the rotor enters the distorted sector, unsteady disturbances in the rotor tip flow are minimized. However, if the rotor meets the casing treatment later in the distorted sector, the disturbances are able to grow, leading to earlier stall inception. Overall, this shows that partial casing treatments can target a distorted flow to give a disproportionately positive impact compared to a full annulus casing treatment. For example, it was found that a 72 deg casing treatment extent could be positioned to achieve over 50% of the full annulus casing treatment stall margin improvement with only 20% of the efficiency penalty.

Loss Assessment of the NASA Source Diagnostic Test Configuration Using URANS With Phase-Lagged Assumption

Abstract This article presents the study of the Source Diagnostic Test fan rig of the NASA Glenn (NASA SDT). Numerical simulations are performed for the three different outlet guide vane (OGV) geometries (baseline, low count, and low noise) and three rotational speeds corresponding to approach, cutback, and sideline operating conditions, respectively. Unsteady Reynolds-averaged Navier–Stokes (URANS) approach is used. The in- and out-duct flow including the nacelle are considered in the numerical simulations, and results are compared against available measurements. Due to the blade count of the fan and OGVs (22 fan blades and either 54 or 26 blades for the OGVs), the simulation can only be reduced to half the full annulus simulation domain using periodic boundary conditions that still represents a significant cost. To alleviate this issue, a URANS with phase-lagged assumption is used. This method allows to perform unsteady simulations on multistage turbomachinery configurations including multiple frequency flows with a reduced computational domain composed of one single-blade passage for each row. The large data storage required by the phase-lagged approach is handled by a compression method based on a proper orthogonal decomposition (POD) replacing the traditional Fourier series decomposition (FSD). This compression method improves the signal spectral content especially at high frequency. Based on the numerical simulations, the flow field is described and used to assess the losses generated in the turbofan architecture based on an entropy approach. The results show different flow topologies for the fan depending on the rotational speed with a leading edge shock at high rotational speed. The fan boundary layer contributes strongly to losses with the majority of the losses being generated close to the leading edge.

Loss assessment of the NASA SDT configuration using LES with phase-lagged assumption

This paper presents the numerical study of the Source Diagnostic Test fan rig of the NASA Glenn (NASA SDT). Large-Eddy Simulations (LES) based on a finite volume approach are performed for the three different Outlet Guide Vane (OGV) geometries (baseline, low-count and low-noise) and three rotational speeds corresponding to approach, cutback and sideline operating conditions respectively. The full stage and nacelle geometries are considered in the numerical simulations, and results are compared to available measurements. The NASA SDT configuration is equipped respectively with 22 fan blades and either 26 of 54 vanes depending on the OGV geometry. The simulation domain could only be reduced to half of the full annulus and would still be a significant cost for the LES. In order to reduce computational cost, an LES with phase-lagged assumption approach is used. This method allows to perform unsteady simulations of multistage turbomachinery configurations including multiple frequency flows with a reduced computational domain composed of one single blade passage for each row. The large data storage required by the phase-lagged approach is handled by a compression method based on a Proper Orthogonal Decomposition replacing the traditional Fourier series decomposition. This compression method improves the signal spectral content especially at high frequency. Based on the numerical simulations, the flow field is described and used to assess the losses generated in the turbofan configuration based on an entropy approach. The results show different flow topologies for the fan depending on the rotational speed with a leading edge shock at high rotational speed. The fan boundary layer contributes strongly to losses with the majority of the losses being generated close to the leading edge for the dissipation due to mean strains and close to the recirculation zone occurring on the suction side for the turbulent kinetic energy production.

Performance Analysis of Non-Interferometry Based Surface Plasmon Resonance Microscopes.

Surface plasmon microscopy has been of interest to the science and engineering community and has been utilized in broad aspects of applications and studies, including biochemical sensing and biomolecular binding kinetics. The benefits of surface plasmon microscopy include label-free detection, high sensitivity, and quantitative measurements. Here, a theoretical framework to analyze and compare several non-interferometric surface plasmon microscopes is proposed. The scope of the study is to (1) identify the strengths and weaknesses in each surface plasmon microscopes reported in the literature; (2) quantify their performance in terms of spatial imaging resolution, imaging contrast, sensitivity, and measurement accuracy for quantitative and non-quantitative imaging modes of the microscopes. Six types of non-interferometric microscopes were included in this study: annulus aperture scanning, half annulus aperture scanning, single-point scanning, double-point scanning, single-point scanning, at 45 degrees azimuthal angle, and double-point scanning at 45 degrees azimuthal angle. For non-quantitative imaging, there is a substantial tradeoff between the image contrast and the spatial resolution. For the quantitative imaging, the half annulus aperture provided the highest sensitivity of 127.058 rad/μm2 RIU−1, followed by the full annulus aperture of 126.318 rad/μm2 RIU−1. There is a clear tradeoff between spatial resolution and sensitivity. The annulus aperture and half annulus aperture had an optimal resolution, sensitivity, and crosstalk compared to the other non-interferometric surface plasmon resonance microscopes. The resolution depends strongly on the propagation length of the surface plasmons rather than the numerical aperture of the objective lens. For imaging and sensing purposes, the recommended microfluidic channel size and protein stamping size for surface plasmon resonance experiments is at least 25 μm for accurate plasmonic measurements.

Assessment of single passage simulations for the aeroelastic stability of a multistage centrifugal compressor

In this paper, the aeroelastic flutter stability of an impeller is investigated numerically with single passage and full annulus URANS simulations carried out on one side on a single stage configuration including a strut and the impeller and on the other side on a multistage configuration in which two additional diffusers are considered downstream. The robustness of single passage simulations relying on the use of multiple frequency phase-lagged boundary conditions is put to the test especially in the case of the multistage configuration for which both blade passage and aeroelastic effects have to be taken into account. Comparisons between single passage simulations on the single and multistage configurations indicate that the impeller-diffuser interaction acts locally on the impeller fluid-structure interface and has only little effect on the generalized aerodynamic forces. Full annulus simulations have also been carried out to assess the accuracy and efficiency of single passage simulations which introduce additional hypotheses of periodicity in the flow. Full annulus simulations are competitive in terms of wall clock since the convergence to the periodic state is faster than with single passage simulations involving the Fourier approximation of the flow field.

Reconstructing Compressor Nonuniform Circumferential Flow Field From Spatially Undersampled Data—Part 2: Practical Application for Experiments

Abstract In the previous part of the paper, a novel method to reconstruct the compressor nonuniform circumferential flow field using spatially undersampled data points is developed. In this part of the paper, the method is applied to two compressor research articles to further demonstrate the potential of the novel method in resolving the important flow features associated with these circumferential nonuniformities. In the first experiment, the static pressure field at the leading edge of a vaned diffuser in a high-speed centrifugal compressor is reconstructed using pressure readings from nine static pressure taps placed on the hub of the diffuser. The magnitude and phase information for the first three dominant wavelets are characterized. Additionally, the method shows significant advantages over the traditional averaging methods for calculating repeatable mean values of the static pressure. While using the multi-wavelet approximation method, the errors in the mean static pressure with one dropout measurement are 70% less than the pitchwise-averaging method. In the second experiment, the full annulus total pressure field downstream of Stator 2 in a three-stage axial compressor is reconstructed from a small segment of data representing 20% coverage of the annulus. Results show very good agreement between the reconstructed total pressure profile and the experiment at a variety of spanwise locations from near hub to near shroud. The features associated with blade row interactions accounting for passage-to-passage variations are resolved in the reconstructed total pressure profile.