Inlet Total Pressure Distortion Research Articles

CFD Analysis and Validation of Axial Compressor Rotor Blade

This research paper employs computational simulations to investigate the consequences of inlet flow distortion, focusing specifically on inlet swirl and total pressure distortion, on the operational characteristics and robustness of an axial transonic compressor. Various permutations of inlet swirl and total pressure distortion are examined, uncovering diverse impacts on the compressor's operational robustness and efficiency. The analysis demonstrates that while co-swirl configurations marginally enhance the operational robustness of the compressor, counter-swirl configurations tend to diminish it. The combination of the effects of individual distortion patterns can provide insights into certain aspects of how a composite pattern of distortion would affect the compressor's efficiency, yet this approach falls short in accurately evaluating the compressor's operational robustness when subjected to a composite distortion pattern.

Effects of Inlet Pressure Distortion on the Performance and Flow Field of a Centrifugal Compressor at Design Condition

The present paper presents experimental results on the effects of inlet total pressure distortion on the performance and flow field of a centrifugal compressor. The total pressure at inlet is artificially distorted by means of a perforated sheet, which is supported by a support mesh. A total of eleven configurations, including clean inlet configuration, are tested. Performance measurements and impeller inlet and exit flow studies are carried out. The distorted inlet flow causes loss of energy, reduction in the maximum flow coefficient and perceptible decrease in the stable operating range of the compressor when compared with undistorted inlet flow. The reduction in stable operating range is the highest for 150° Distortion and the least for 4x22.5° Distortion. In radial distortion group, tip distortion has more detrimental effect on the operating range, energy coefficient and efficiency. The mass averaged total pressure at the impeller exit is the highest for clean inlet and the lowest for Hub+90° Distortion. The circumferential distortion has more adverse effect than the radial distortion on the flow parameters. The combined radial and circumferential distortion (Hub+90°) has the worst effect on the mass averaged flow parameters. As the Distortion Index (DI) increases, the mass averaged total pressure at exit stations decreases. Distortion sector angle of 60° having the lowest total pressure is found to be the critical sector for circumferential distortion configurations. As the Distortion Correlation parameter, DC(60) increases, the mass averaged total pressure for circumferential distortion configuration decreases.

Effects of Stage Loading on Performance and Flow Field of a Centrifugal Compressor With Inlet Pressure Distortion

The paper presents experimental results on the effects of inlet total pressure distortion on the performance and flow field of a centrifugal compressor. The total pressure at the inlet is artificially distorted by means of a perforated sheet, which is supported by a support mesh. A total of eleven configurations, including clean inlet configuration, are tested. Performance measurements and impeller inlet and exit flow studies at three flow coefficients, one near design flow coefficient, one below design flow coefficient and one above design flow coefficient, are carried out. The present paper presents and discusses results at the three flow coefficients and the effects of stage loading on the distortion effects are presented. A new parameter, Distortion Index (DI) is introduced. As DI increases, the mass averaged total pressure at exit stations decreases. Distortion sector angle of 60° having the lowest total pressure is found to be the critical sector for circumferential distortion configurations. As the Distortion Correlation parameter, DC (60) increases, the mass averaged total pressure for circumferential distortion configuration decreases, except in the case of low flow coefficient where the mass averaged total pressure is nearly constant. DC (60) also increases with sector angle. The static pressure normalized with static pressure for clean inlet decreases as the distortion sector angle is increased. Distortion attenuates the static pressure as the flow passes through the vaneless diffuser. The attenuation increases with the distortion sector angle.

Analysis of the Unsteady Flow Field at Stall Conditions for a Low-Speed Low-Pressure Ratio Axial Fan With Full-Annulus Simulation

Abstract The present work analyzes the three-dimensional unsteady flow structure of a low-speed low-pressure ratio axial fan stage operating at stall conditions. Unsteady numerical simulations were conducted with the elsA software, which solves the unsteady Reynolds-averaged Navier–Stokes (URANS) equations on a full-annulus calculation domain for the rotor and the stator. The analyses will focus on a near stall operating point, the stall inception mechanisms, and the final rotating stall pattern. Numerical simulations are confronted with experimental measurements from high-frequency Kulite transducers mounted on the casing near the rotor blades and the stator blades. The results show that boundary layer separation occurs on the stator blade close to the hub region at near stall conditions. The reverse flow interacts with the main flow to form a vortex shedding phenomenon. The frequency is not clocked with the blade passage frequency, requiring unsteady full-annulus methodology even for a stable operating point. It is also shown that the final state of the rotating stall region is composed of a single part-span cell that occupies three-quarters of the circumference and rotates at 60% of the shaft speed. This pattern agrees with the experimental measurements in terms of both the number of cells and the rotational speed. Finally, the paper demonstrates that without inlet distortion, the mechanism of the onset of the rotating stall is incorrectly predicted by the simulation. However, when a total pressure inlet distortion is applied, the computational fluid dynamics produces the same inception mechanism that is observed experimentally.

Influence of inlet distortion on the wear of aero-compressor blades

The turboshaft engine is the main power source of a helicopter. Under airborne conditions, inlet distortion inevitably occurs at the entrance of the compressor because of factors such as airflow and environmental turbulence. The change in the internal flow field alters the erosive wear law of compressor blades caused by sand particles. The influence of total pressure inlet distortion on the erosive wear of compressor blades is studied innovatively in this paper. A new parameterization method of the erosive wear model is proposed through particles velocity test and erosive wear experiment. The rotor blades of pre-1.5 stage compressor are used as the research object. A wear analysis model of the full-port compressor blades is established, which can take into account the inlet distortion. The influence mechanism of typical inlet distortion angles on the erosive wear distribution characteristics and wear degree of the rotor blades is revealed. The results show that inlet distortion significantly alters the wear characteristics of the rotor blades. Specifically, owing to the pressure difference and suction effect of the high-speed rotor blade, the sand particles in the distortion-affected area exhibit aggregation, re-aggregation, and transfer phenomena, which leads to considerable differences in the distribution characteristics and degree of the erosive wear between the rotor blades’ entering the distortion area and their exiting the distortion area. When the inlet distortion angle is expanded from 30° to 90°, compared with the rotor blade in the non-distortion-affected area, the wear degrees of the rotor blade affected by the inlet distortion increase by 32.5–68.7% at the leading edge and by 51.3–57% at the middle position. Moreover, when the inlet distortion angle increases, the maximum wear rate concentration values at the leading edge of the blade with the most serious wear and second most serious wear in the entire circumference of the rotor blade decrease by 4.26% and 27.3%, respectively.

Effects of tip air injection on the aerodynamic stability of an axial flow compressor with total pressure distortion

The compressor is a critical component that determines the aerodynamic stability of an aero-engine. Total pressure inlet distortion decreases the thrust and shrinks the stability margin, thus inducing severe performance degradation or even flameout. Generally, tip air injection is used to reduce the adverse influence of total pressure inlet distortion on the aerodynamic stability. In the present work, an experimental investigation on the effects of tip air injection on the stability of a two-stage low-speed axial compressor with total pressure inlet distortion was carried out. A flat baffle generated the total pressure distortion at the inlet of the compressor. The stall margin of the compressor was reduced significantly by the total pressure distortion. When the dimensionless insertion depth of the flat baffle was 0.45, the stall margin decreased to 11.4%. Under the total pressure inlet distortion, tip air injection effectively improved the distortion resistance capability of the compressor. The circumferential layout of the nozzle played a critical role in the stability expansion effect of tip air injection under the inlet flow condition of the total pressure distortion. The modal wave disturbance was likely to occur in the distortion-affected region (the low-pressure region and the mixing region). Tip air injection did not inhibit the generation of the modal wave but restrained the development of the modal wave into the stall cell. It improved the low-speed compressor’s tolerance to the modal wave and allowed a higher amplitude modal wave to occur.

Assessment of the Influence of Boundary Layer Ingestion (BLI) on the Axial Fan

Boundary Layer Ingestion (BLI) is a technology which is effective to reduce fuel consumption of future commercial aircrafts. However, the inlet distortion caused by BLI has a negative on the engine performance. This paper investigates the effect of the inlet total pressure distortion on the aerodynamic performance of the fan rotor. Unsteady numerical analysis was conducted on JAXA TechClean Fan with circumferential distortion caused by BLI. As the result, under the high distorted condition, the peak efficiency of the rotor decreases by 0.622 points. It is found that the main factor of the efficiency reduction is tip-wall passage vortex.

Experimental and numerical investigation on total pressure distortion generated by a rectangular movable flat baffle

A distortion generator equipped with a motor-activated movable flat baffle was installed just upstream of a rectangular plenum entrance to investigate the effects of inlet total pressure distortion on the stability and performance of an auxiliary power unit (APU). Experiments and numerical simulations on a direct connect scale inlet model of the APU were carried out to obtain a quantitative relationship between the insertion depth of the flat baffle in the flow stream and the total pressure distortion intensity and region. In the experiments, the blocking coefficient and total pressure distortion coefficient were controlled by adjusting the insertion depth of the flat baffle and the mass flow. In the simulations, detailed flow field was analyzed based on the detached-eddy simulation (DES) method. The results show that the pressure distribution of the distorted flow on the aerodynamic interface plane (AIP) can be divided into a high-pressure region, a transition region, and a low-pressure region. The area affected by the distorted flow was larger than the inserting area of the flat baffle. That area was more related to the relative blocking coefficient, and less affected by the mass flow. The total pressure distortion coefficient had a linear relationship with the mass flow rate and is positively correlated with the relative blocking coefficient. As the relative blocking coefficient increased to a certain value, an exponential growth in the total pressure distortion coefficient occurred, and consequently, the flow field distortion was intensified. In the flow field, a pair of corner vortices were formed at the corner between the flat baffle and the bottom wall of the inlet pipe, and a large separation zone was formed behind the flat baffle and exhibits certain unsteady characteristics.

Prediction of engine total pressure distortion in improved cascaded forward network

In order to solve the problem of the result delay in the real-time calculation of engine inlet total pressure distortion index in flight test, the multi-step prediction of engine inlet total pressure distortion index in flight test is carried out. The average prediction error of traditional cascade forward neural network prediction model is higher than traditional autoregressive integrated moving average model. An improved algorithm is proposed. By establishing a time series dynamic level model, the time series of engine inlet total pressure distortion index is divided into low dynamic series and high dynamic series by using particle swarm optimization algorithm. The cascade forward neural network prediction model is used for training and prediction respectively. The results show that the average prediction error and maximum prediction error of the improved algorithm are reduced by 3.90%, 10.66% and 3.29% and 1.38% respectively compared with autoregressive integrated moving average model and traditional cascaded feedforward neural network.

Numerical Research on Inlet Total Pressure Distortion in a Transonic Compressor with Non-Axisymmetric Stator Clearance

Inlet total pressure distortion has great adverse effects on the aero-engine performance. The distorted flow passes through the compressor and becomes non-uniform in the downstream blade rows. Different from previous studies based on the assumption of circumferential uniformity, this study aims to improve circumferential non-uniform flow with the non-axisymmetric structure. Non-axisymmetric stator clearance was adopted to resolve the effects of non-uniform flow caused by inlet total pressure distortion in this paper. The 9 stators with tip clearance were installed in the distorted region and the flow field structure and performance under different operating conditions was studied. The study finds that the non-axisymmetric compressor with 9 tip clearance stators can ensure compressor efficiency while improving compressor stability margin. What’s more, the separation range and strength in the distorted region can be reduced significantly and the anti-distortion capability of compressor can be enhanced.

Optimization of a transonic axial-flow compressor under inlet total pressure distortion to enhance aerodynamic performance

In a boundary layer ingestion (BLI) propulsion system, the compressor rotor operates under the condition of a severe inflow distortion because the boundary layers on the surface of aircraft have been ingested. In this situation, the low-momentum fluid would inevitably raise the blade load and aggravate the local aerodynamic losses, thus severely deteriorating the compressor overall aerodynamic performance. The paper has applied a sweep optimization design based on surrogate model and genetic algorithm to a transonic axial-flow compressor at the condition of total pressure distorted inflow. The optimization objective is to improve the compressor efficiency. The sweep design has been implemented through controlling the shape of the blade stacking line and the time-saving steady numerical simulations are mainly utilized in the optimization process. Then, the full-annulus unsteady CFD calculations have been conducted for validating the optimized design and further analyzing the flow mechanisms of performance improvements. The computational results show that the inflow distortion has led to non-uniform distributions of the tip leakage loss and the hub flow separation loss along the circumferential direction. It is found that the rotor blade has suffered from the highest tip leakage losses and hub separation losses during the blade leaving the distorted region. After rotor sweep optimization, both tip leakage and hub separation losses have been notably reduced along the full circumferential direction. Moreover, for different pitchwise locations, the optimized rotor sweep design has achieved more aerodynamic performance improvements in the process of the rotor blade leaving the distorted region than those in the processes of the rotor blade moving out of, moving towards and entering the distorted region. Overall, the optimized design performs better than the baseline design in terms of adiabatic efficiency in the entire operating range. In the meanwhile, the total pressure ratio has been maintained. Specifically, an approximately 0.5% increment of efficiency has been achieved by the optimized design at the optimization point. For this paper, the major purpose is to offer useful guidelines for fan/compressor design strategy to achieve high aerodynamic performances under distorted inflow working condition.

A mixed-fidelity computational model of aero engine for inlet distortion

This paper presents a mixed-fidelity model to investigate the effect of complex inlet distortion on aero gas turbine engines. The approach is developed by integrating a three-dimensional body force model of multistage compressors and a classical two-dimensional parallel engine model. The internal flow field of a high bypass ratio turbofan engine under different forms of total pressure distortion is simulated by the model. The simulations are discussed in depth to reveal the features of the distorted flow field in the compression components under the whole engine environment. The relationship between the overall performance of the engine and the intensity of inlet total pressure distortion is investigated quantitatively. It is evident that the mixed-fidelity model has the capability to quantitatively evaluate the effect of complex inlet distortion on the performance and the internal flow field of engines with low computational costs.

Experimental study of a controlled variable double-baffle distortion generator engine test rig

In order to explore the total-pressure distortion test assessment method for a turbofan engine, a Controlled Variable Double-Baffle Distortion Generator (CVDBDG) with a horizontal symmetry moving form was developed, which can adjust the steady-state and time–variant distortion separately in real time. The inlet total-pressure distortion test was conducted on an afterburner turbofan engine. The distortion parameters of CVDBDG and the instability characteristics of the engine were measured. The experimental data were modeled and analyzed by using back propagation artificial neural networks, and the work envelope of CVDBDG was obtained. Based on the analysis of the data on the engine’s instability, the properties of CVDBDG used for the stability assessment were preliminarily evaluated. The results show that CVDBDG can simulate both steady-state and time–variant distortions simultaneously in a range determined by three envelopes. Under the condition of symmetric double baffles, a critical depth of insertion exists, beyond which the symmetric baffles will generate an asymmetric flow field. In the case of double baffles, compared to a single baffle, the engine exhibited different instability characteristics. Based on CVDBDG, it is expected that more efficient engine stability and durability assessment methods can be developed.

Experimental Investigation of Inlet Distortion Effect on Performance of a Micro Gas Turbine

An experimental study has been carried out to determine how inlet total-pressure distortion affects the performance of a micro gas turbine. An inlet simulator is designed and developed to produce and measure distortion patterns at the inlet to the gas turbine. An air jet distortion generator (AJDG) is used to produce nonuniform flow patterns and total pressure probes are installed to measure steady-state total-pressure distribution at the inlet. A set of wind tunnel tests have been performed to confirm the fidelity of distortion generator and measuring devices. Tests are carried out with the gas turbine exposed to inlet flow with 60 deg, 120 deg, and 180 deg circumferential distortion patterns with different distortion intensities. The performance of the gas turbine has been measured and compared with that of clean inlet flow case. Results indicate that the gas turbine performance can be affected significantly facing with intense inlet distortions.

Assessing the Impact of the Inlet Total Pressure Distortion on the Turbofan Thrust

The paper considers techniques to develop a mathematical model using a method of «parallel compressors». The model is intended to estimate the impact of the air inlet distortion on the primary parameters of the aero-engine.  The paper presents rated estimation results in the context of twin spool turbofan design for two typical cruiser modes of flight of the supersonic passenger jet. In estimation the base values σbase and the average values of the inlet ram recovery σave remained invariable. Thus, parametrical calculations were performed for each chosen relative value of the area of low-pressure region.The paper shows that an impact degree of the inlet distortion on the engine thrust for two modes under consideration is essentially different. In other words, if in the subsonic mode the impact assessment can be confined only to taking into account the influence of decreasing average values of the inlet total pressure, the use of such an assumption in the supersonic cruiser mode may result in considerable errors.With invariable values of the pressure recovery factor at the engine intake, which correspond to the speed of flight for a typical air inlet of external compression σbase, and average value σave, a parameter Δσuneven  has the main effect on the engine thrust, and degree of this effect essentially depends on a difference between σave and σbase values.

Numerical research on inlet total pressure distortion in a transonic compressor with non-axisymmetric stator

The adverse impacts of non-uniform inlet flow have been the focus for several decades with the increase of the operating range of engines. A deep understanding of the flow mechanism of distortion passing through a compressor is needed urgently and the improvement of the compressor performance becomes more and more important. In this paper, a non-axisymmetric stator is presented with significant non-axisymmetric characteristics in a transonic compressor to investigate compressor performance and flow field effects. A time-dependent three-dimensional Reynolds-averaged Navier-Stokes equation composed in ‘Fluent Software Pack’ is validated and used to perform the simulations. The flow fields with distorted inlet are obtained and the effects of original stator and non-axisymmetric stator in a transonic compressor are compared. The results are discussed in terms of the effects of non-axisymmetric stator on compressor performance, blockage of flow passage, rotor and stator. The results show that the non-axisymmetric stator influences not only the stator flow field but also the rotor flow field, so the efficiency and total pressure ratio are improved correspondingly.

Influence of SPS casing treatment on axial flow compressor subjected to radial pressure distortion

The present work is about the stall margin enhancement ability of a kind of stall precursor-suppressed (SPS) casing treatment when fan/compressor suffers from a radial total pressure inlet distortion. Experimental researches are conducted on a low-speed compressor with and without SPS casing treatment under radial distorted inlet flow of different levels as well as uniform inlet flow. The distorted flow fields of different levels are generated by annular distortion flow generators of different heights. The characteristic curves under these conditions are measured and analyzed. The results show that the radial inlet distortion could cause a stall margin loss from 2% to 30% under different distorted levels. The SPS casing treatment could remedy this stall margin loss under small distortion level and only partly make up the stall margin loss caused by distortion in large level without leading to perceptible additional efficiency loss and obvious change of characteristic curves. The pre-stall behavior of the compressor is investigated to reveal the mechanism of this stall margin improvement ability of the SPS casing treatment. The results do show that this casing treatment delays the occurrence of rotating stall by weakening the pressure perturbations and suppressing the nonlinear amplification of the stall precursor waves in the compression system.

Numerical investigation of effect of inlet swirl and total-pressure distortion on performance and stability of an axial transonic compressor

This paper represents numerical simulation of flow inside an axial transonic compressor subject to inlet flow distortion, to evaluate its effect on compressor performance and stability. Two types of inlet distortion, namely inlet swirl and total pressure distortion are investigated. To study the effect of combined distortion patterns, different combinations of inlet swirl and total pressure distortion are also studied. Results for cases with total pressure distortion indicate that hub radial distortion improves stability range of the compressor while tip radial distortion deteriorates it. An explanation for this observation is presented based on redistribution of flow parameters caused by distortion and the way it interacts with stall inception mechanisms in a transonic axial compressor. Results also show that while co-swirl patterns slightly improve stability range of the compressor, counter-swirl patterns diminish it. Study of combined distortion cases reveals that superimposition of effects of each individual pattern could predict the effect of a combined pattern on compressor’s performance within an accuracy of 1%. However, it is unable to predict the associated effect on compressor’s stability.

Effect of inlet distortion on the performance of axial transonic contra-rotating compressor

The present paper numerically conducted full-annulus investigation on the effects of circumferential total pressure inlet distortion on the performance and flow field of the axial transonic counter-rotating compressor. Results reveal that the inlet distortion both deteriorates the performance of the upstream and downstream rotors resulting in reduction of total pressure ratio, efficiency and stall margin of the transonic contra-rotating compressor. Regarding the development of distortion inside compressor, the downstream rotor reinforces the air-flow mixing effects and, thus, attenuates the distortion intensity significantly. Under the distorted inflow conditions, the detached shockwave at the leading edge of downstream rotor interacts with the tip leakage flow and causes the blockage of the blades passage, which is one important reason for the transonic contra-rotating compressor stall.

Bleed-Induced Distortion in Axial Compressors

In this paper, the influence of nonuniform bleed extraction on the stability of an axial flow compressor is quantified. Nonuniformity can be caused by several geometric factors (for example, plenum chamber size or number of off-take ducts), and a range of configurations is examined experimentally in a single stage compressor. It is shown that nonuniform bleed leads to a circumferential distribution of flow coefficient and swirl angle at inlet to the downstream stage. The resultant distribution of rotor incidence causes stall to occur at a higher flow coefficient than if the same total bleed rate had been extracted uniformly around the circumference. A connection is made between the analysis of nonuniform bleed extraction and the familiar DCθ criterion used to characterize inlet total pressure distortion. The loss of operating range caused by the nonuniform inlet flow correlates with the peak sector-averaged bleed nonuniformity for all the bleed configurations tested.