Adiabatic Efficiency Research Articles

Experimental research on influence of interstage oil injection in a dual-motor-driven two-stage screw air compressor

Energy saving for compressed air production is of great importance all over the world due to its great energy cost and the indispensable utility of compressed air in the industrial sector. In recent years, oil-injected two-stage screw air compressors, especially driven by dual motors, are becoming a promising alternative to other compression equipment for their extraordinary efficiency and high reliability in a wide range of applications. However, two-stage screw air compressors have a different and complicated oil injection system compared with single-stage compressors. And an appropriate setting of oil injection parameters might result in a considerable performance improvement of the compressor. In this paper, a dual-motor-driven two-stage oil-injected screw compressor was tested under the actual service conditions and control modes experimentally. The effect of interstage oil injection on the compressor performance was studied in detail by varying the interstage oil injection flow rate and rotating speed. The performance parameters such as specific power, volumetric efficiency and adiabatic efficiency were calculated from the measured data. The results indicate that an excess amount of oil exists in the second stage compressor and reducing the oil moderately could decrease the power consumption of the two-stage air compressor unit.

Waste heat recovery of an UAV propulsion system based on PEM fuel cell by a novel transcritical CO2 - LNG hybrid cycle; thermodynamic and multiple linear regression analyses

ABSTRACT In this work, the performance of a proposed hybrid system consisting of a transcritical carbon dioxide-natural gas cycle for recovery of waste heat produced in a proton exchange membrane fuel cell has been evaluated. The effects of different parameters on proposed cycle performance have been investigated thermodynamically. This hybrid system can convert fuel cell’s released heat to electricity, augment production power, convert natural gas from the liquid phase to gas, and appropriate use of heat energy from consumed water. Results show that increasing the cell’s power by about 52% leads to a growth in overall efficiency of transcritical power cycle. Also, in a constant fuel cell power, the higher carbon dioxide mass flow rate tends to lower carbon dioxide turbine output power, and consequently, the efficiency of transcritical power cycle decreases. Changes in waste heat recovery efficiency among models with different values of CO2 mass flow rate increases very slightly as fuel cell power augments. For instance, at lowest examined fuel cell power (790 kW), growth in CO2 mass flow rate by about 26% causes a decline in cycle efficiency by about 6.35%. By increasing fuel cell thermal efficiency by 30%, improvement in overall cycle efficiency rises significantly by about 109.92%. On the other hand, augmentation in fuel cell thermal efficiency by about 30% leads to a very slight growth in improvement of the waste heat recovery efficiency by about 2.67%. Moreover, by increasing adiabatic efficiency of turbines by about 21.43%, improvement in waste heat recovery efficiency declines by about 10.75%. Increasing carbon dioxide turbine’s inlet temperature and reducing carbon dioxide mass flow rate have beneficial effects on efficiency of carbon dioxide transcritical power cycle. Eventually, two high-accuracy models have been presented for output power and efficiency of proposed hybrid cycle using multiple linear regression.

Aerodynamic Optimization of Transonic Rotor Using Radial Basis Function Based Deformation and Data-Driven Differential Evolution Optimizer

The complicated flow conditions and massive design parameters bring two main difficulties to the aerodynamic optimization of axial compressors: expensive evaluations and numerous optimization variables. To address these challenges, this paper establishes a novel fast aerodynamic optimization platform for axial compressors, consisting of a radial basic function (RBF)-based blade parameterization method, a data-driven differential evolution optimizer, and a computational fluid dynamic (CFD) solver. As a versatile interpolation method, RBF is used as the shape parameterization and deformation technique to reduce optimization variables. Aiming to acquire competitive solutions in limited steps, a data-driven evolution optimizer is developed, named the pre-screen surrogate model assistant differential evolution (pre-SADE) optimizer. Different from most surrogate model-assisted evolutionary algorithms, surrogate models in pre-SADE are used to screen the samples, rather than directly estimate them, in each generation to reduce expensive evaluations. The polynomial regression model, Kriging model, and RBF model are integrated in the surrogate model to improve the accuracy. To further save optimization time, the optimizer also integrates parallel task management programs. The aerodynamic optimization of a transonic rotor (NASA Rotor 37) is performed as the validation of the platform. A differential evolution (DE) optimizer and another surrogate model-assisted algorithm, committee-based active learning for surrogate model assisted particle swarm optimization (CAL-SAPSO), are introduced for the comparison runs. After optimization, the adiabatic efficiency, total pressure ratio, and surge margin are, respectively, increased by 1.47%, 1.0%, and 0.79% compared to the initial rotor. In the same limited steps, pre-SADE gets a 0.57% and 0.51% higher rotor adiabatic efficiency than DE and CAL-SAPSO, respectively. With the help of parallel techniques, pre-SADE and DE save half the optimization time compared to CAL-SAPSO. The results verify the effectiveness and the rapidity of the fast aerodynamic optimization platform.

Multi-Objective Optimization of Aerodynamic Performance for a Small Single-Stage Turbine

In order to improve the performance of single-stage turbines, blade profiling optimization was conducted for the guide vane and rotor under design condition. Support vector regression (SVR) and non-dominated sorting genetic algorithm-II (NSGA-II) were used to execute the optimization, with the objective of maximizing the efficiency and total pressure ratio of single-stage turbines. The gas turbine chosen for the initial study was the KJ66, which is one of the most robust and primitive small gas turbine designs available. The influence mechanism of the stator and rotor profiling on flow field and performance was discussed. The results revealed that compared with the prototype, the adiabatic efficiency increased by 5.95% and the total pressure ratio increased by 0.9%. Furthermore, the matching of flows between the stator blade and rotor blade obviously improved. The optimized guide vane suppressed the flow separation by increasing the leading edge and improving the distribution of the inlet angle of attack. The load distribution of rotors with a 50% spanwise position changed from the original "C" loaded to post-loaded. The leading load obviously decreased, and the angle of attack was smaller than that of the prototype, which effectively weakened the flow separation at the leading edge of the rotor. Compared with the original rotor, the higher lean angle and pressure ratio of the turbine stage also improved. However, the leakage loss near the shroud of the rotor increased, which led to decreased efficiency.

Numerical Investigations of a Non-Uniform Stator Dihedral Design Strategy for a Boundary Layer Ingestion (BLI) Fan

A distributed propulsion system has the advantage of saving 5–15% fuel burn through ingesting the fuselage boundary layer of an aircraft by fan or compressor. However, due to boundary layer ingestion (BLI), the fan stage will continuously operate under serious inlet distortion. This will lead to a circumferentially non-uniform flow separation distribution on the stator blade suction surface along the annulus, which significantly decreases the fan’s adiabatic efficiency. To solve this problem, a non-uniform stator dihedral design strategy has been developed to explore its potential of improving BLI fan performance. First, the stator full-annulus blade passages were divided into blade dihedral design regions and baseline design regions on the basis of the additional aerodynamic loss distributions caused by BLI inlet distortion. Then, to find the appropriate dihedral design parameters, the full-annulus BLI fan was discretized into several portions according to the rotor blade number and the dihedral design parameter investigations for dihedral depth and dihedral angle were conducted at the portion with the largest inflow distortion through a single-blade-passage computational model. The optimal combinational dihedral design parameter (dihedral depth 0.3, dihedral angle 6 deg) was applied to the blade passages with notable flow loss which were mainly located in the annulus positions from −120 to 60 degrees suffering from inlet distortion, while the blades in the low-loss annulus locations were unchanged. In this way, a non-uniform stator dihedral design scheme was achieved. In the end, the effectiveness of the non-uniform stator dihedral design was validated by analyzing the internal flow fields of the BLI fan. The results show that the stator dihedral design in distorted regions can increase the inlet axial velocity and reduce the aerodynamic load near the blade trailing edge, which are beneficial for suppressing the flow separations and reducing aerodynamic loss. Specifically, compared with the baseline design, the non-uniform stator dihedral design has achieved a reduction of aerodynamic loss of about 7.7%. The fan stage has presented an improvement of adiabatic efficiency of about 0.48% at the redesigned point without sacrificing the total pressure ratio. In the entire operating range, the redesigned fan has also shown a higher adiabatic efficiency than the baseline design with no reduction of the total pressure ratio, which provides a probable guideline for future BLI distortion-tolerant fan design.

Full-envelope acceleration control method of turbofan engine based on variable geometry compound adjustment

In order to reduce the influence of compressor stability margin constraint and enhance the full envelope acceleration performance of turbofan engine, a full envelope acceleration control method of turbofan engine based on variable compressor intermediate stage bleed and adjustable low-pressure (LP) turbine guide vane is proposed. Firstly, a high confidence turbofan engine component level model is established, which is capable of modifying the compressor adiabatic efficiency caused by variable intermediate stage bleed and adjusting the LP turbine guide vane adaptively. Then, the main combustion chamber fuel flow, compressor intermediate stage bleed flow and LP turbine guide vane angle are selected as the optimization variables. Meanwhile, the PSO-SQP (Particle Swarm Optimization-Sequential Quadratic Programming) method is adopted to obtain the acceleration compound control schedule. Finally, the full envelope compound adjustment acceleration control plan is available based on the similarity conversion theory of consistent engine inlet total temperature. The simulation results demonstrate that the acceleration control method based on variable geometry compound adjustment can significantly promote engine acceleration performance in the left flight envelope, shorten the maximum acceleration time by more than 50%, and dramatically improve the stability margin of the compressor during acceleration phase. In the right flight envelope, the variation law of the minimum surge margin during acceleration procedure almost coincides with the engine inlet isotherm. It proves that the method proposed possesses good applicability and feasibility in the full envelope.

Research on 3D Design of High-Load Counter-Rotating Compressor Based on Aerodynamic Optimization and CFD Coupling Method

In view of the flow instability problem caused by the strong shock wave and secondary flow in the channel of the high-load counter-rotating compressor, this paper adopts the design method of coupling aerodynamic optimization technology and CFD and establishes a three-dimensional aerodynamic optimization design platform for the blade channel based on an artificial neural network and genetic algorithm. The aerodynamic optimization design and internal flow-field diagnosis of a high-load counter-rotating compressor with a 1/2 + 1 aerodynamic configuration are carried out. The research indicates that the optimized blade channel can drive and adjust the flow better, and the expected supercharging purpose and efficient energy conversion process are achieved by controlling the intensity of the shock wave and secondary flow in the channel. The total pressure ratio at the design point of the compressor exceeds 2.9, the adiabatic efficiency reaches 87%, and the aerodynamic performance is excellent at the off-design condition, which is on the advanced design level of the same type of axial compressor. The established aerodynamic optimization design platform has important practical engineering applications for the development of high thrust-to-weight ratio aero-engine compression systems.

Experimental investigation on helium valved linear compressors with different active offsets

Piston offset in the helium valved linear compressor (VLC) deteriorates the cooling performance and operating life of Joule-Thomson (JT) refrigerators, especially in aerospace applications. To clarify the influence of piston offset on the VLC performance, an experimental investigation was carried out. Based on a laboratory developed VLC, the effects of piston offset on the suction and discharge pressure, pressure ratio, mass flow rate, motor efficiency, volume efficiency and adiabatic efficiency are systematically studied under two piston motion states (fixed piston displacements and full strokes). The results show that when the piston stroke is constant, the offset will increase the discharge pressure and mass flow rate, and decrease suction pressure. When the piston moves at full stroke, the offset will reduce the discharge pressure and mass flow rate, and increase suction pressure. Regardless of the working conditions, the existence of offset will affect the efficiency of the compressor, hence increasing the power consumption.

AERODYNAMIC PERFORMANCES OF A SINGLE-STAGE TRANSONIC AXIAL COMPRESSOR USING AN INCLINED CASING GROOVE

ABSTRACT This paper presents the effect of an inclined casing groove on the aerodynamic performance of a single-stage transonic axial compressor, NASA Stage 38, using three-dimensional Reynold-averaged-Navier-Stokes equations with the k-ε turbulence model. The research was carried out to examine the effects of four casing groove parameters: angle, width, depth, and location. Validation of a numerical model for a single-stage transonic axial compressor was conducted to evaluate the computational fluid dynamics method. Most of the simulations showed positive results with an increase in stall margin, adiabatic efficiency, and total pressure ratio, in which the maximum stall margin, adiabatic efficiency, and total pressure ratio can be raised by 87.09%, 0.13%, and 1.57%, respectively, as compared to the smooth casing case.

Critical endwall blockage attenuation-based automatic optimization of casing treatment design for transonic axial flow compressor

Tip leakage flow (TLF) and endwall blockage formation are closely associated with the inception of rotating stall in tip-critical transonic axial flow compressor (AFC) rotors. The present study proposes a critical endwall blockage attenuation-based optimization method of casing treatment design for transonic AFC. First, a well-defined endwall blockage quantification method is developed to reveal the development of the endwall blockage at near-stall, and two blockages of the maximum endwall blockage and the near trailing edge endwall blockage are identified as the critical values for stall inception. Next, both adiabatic efficiency at design point and critical endwall blockages at near-stall are used as optimization objectives in an adaptive surrogate-based optimization procedure to deal with the design of nonuniform multiple grooves casing treatment (NMGCT) for the NASA Stage 35, thus avoiding the extremely time-consuming simulations of compressor performance curves. The analysis of the representative optimal designs demonstrates that the NMGCT can act on the tip flow field in a variety of ways, e.g. lowering the maximum endwall blockage can suppress the intensity of the TLF-shock wave interaction, or lowering the near trailing edge endwall blockage is able to control both the trajectory of the TLF closer to the suction surface and the thickness of the suction surface boundary layer. These achievements can help improving the stall margin of the tip-critical transonic AFC. However, in the case of AFC similar to Stage 35, reducing the maximum endwall blockage can achieve greater stall margin improvement and smaller peak efficiency penalty.

Research on infilling strategy in surrogate-aided optimization for axial compressor blades

The hybrid Voronoi-Latin Hypercube Sampling (Voronoi-LHS) method is proposed for the surrogate-aided optimization of the axial compressor blades. The hybrid method is first applied to the fundamental test cases. The analytical results show that, compared with the Voronoi and LHS strategies, the hybrid method generally improves the robustness and convergency. Then the multi-objective genetic algorithm (MOGA) in conjunction with the artificial neural network (ANN) is applied to optimize the aerodynamic performance of an axial compressor rotor. Before the optimization process, the hybrid Voronoi-LHS sample infilling method is employed to refine the ANN surrogate model. Considering the typical intake distortion, the sweep and lean distributions of this rotor are optimized to pursue the maximum total pressure ratio and adiabatic efficiency. The results show that the optimization significantly improves the pressure ratio, efficiency and surge margin of the compressor with low computing cost.

Influence of clearance height on performance of water-lubricated single-screw air compressor

Addressing the growing need of oil-free compressed air has drawn the development of the water-lubricated single-screw air compressor (WLSSAC). For WLSSAC, clearance height is an important design parameters relevant to compressor performance. In this study, a thermodynamic model of the working process of a WLSSAC is established in consideration of the viscous friction power consumption and mass transfer between water and wet air. The theoretical model is simulated by Matlab software, and the accuracy of the theoretical model is verified by experimental results. The influence of each leakage path on the rate of leakage flow, viscous friction power consumption, and efficiency are analyzed using the proposed model. The effects of clearance height on adiabatic efficiency, volumetric efficiency, and specific power are calculated and discussed. Results show that the viscous friction power consumption caused by wall velocity decreases gradually with the increase in clearance height, while the viscous friction power consumption caused by pressure difference increases gradually in each leakage path. With the increase in clearance height, the rate of leakage flow increases gradually, the adiabatic and volumetric efficiencies decrease gradually, and the specific power increases gradually. The clearance height is recommended to be 0.03–0.06 mm for the design of WLSSAC.

Casing treatment using oblique slots for a single-stage transonic axial compressor

Abstract This work presents a casing treatment design using oblique slots to improve the stability of a single-stage transonic axial compressor, NASA Stage 37. 36 individual slots are distributed uniformly on the rotors’ shroud. Numerical simulations were performed to find the effects of the casing treatment on the compressor’s stability as well as the adiabatic efficiency and pressure ratio. The oblique slots were found to effectively alleviate the negative influence of the tip leakage flow on the compressor’s performance, thus increasing the stall margin without significant sacrifices in efficiency and pressure rise. Three geometric parameters of the casing treatment were examined. The results showed that, with the oblique slots, the highest stall margin reached 13.59% - a 37% increase from the smooth casing case, with about 0.17% and 0.09% reductions in the peak efficiency and maximum pressure ratio, respectively.

Automatic Design Optimization of a Transonic Compressor Rotor for Improving Aeroacoustic and Aerodynamic Performance

Abstract Multidisciplinary design optimization of transonic rotors, especially focusing on improving the aeroacoustic performance, has received little attention. This article employs an automatic design optimization approach to reduce the shock-associated tone noise and to increase the adiabatic efficiency of NASA rotor 37. An evolutionary algorithm for multi-objective trade-offs searching is used to obtain the noise-efficiency Pareto front. Three-dimensional Reynolds-averaged Navier–Stokes simulations are directly applied in the optimization loop for aeroacoustic and aerodynamic analysis. Significant noise reduction and efficiency increase have been achieved. Two optimized designs corresponding to the endpoints of the Pareto front have been analyzed comprehensively from the perspectives of rotor characteristics, noise performance, and flow and acoustic fields, to understand the mechanisms responsible for performance improvements.

Evaluation of non-uniform water film behavior on the performance of an axial aeroengine compressor

Abstract Two-phase flow CFD simulations are conducted to investigate the performance characteristics of an axial flow compressor operating under an excessive amount of water intrusion exposed to a rainy atmosphere. In order to obtain a certain compressor behavior under the water film irregularities, the blade surfaces are divided into a small number of regions depending on the film thickness. In addition, to examine the droplets interaction with rotating blades and splashing process, a self-compiled FORTRAN code has been used. The results showed that uneven roughening of blade surfaces is the effective technique to identify the deterioration in performance with irregular wall film thickness. The overall adiabatic efficiency and pressure ratio are decreased by 6% and 0.88 at a 5% water injection rate using an irregular rough wall model. Moreover, the increased wall film has adversely influenced the internal flow mechanism, which resulted the increase of passage shock intensity, entropy level within the main flow zone as well as the blade surface temperature.

Investigation on shock-induced separation loss mitigation method considering radial equilibrium in a transonic compressor rotor

AbstractA shock-induced separation loss reduction method, using local blade suction surface shape modification (smooth ramp structure) with constant adverse pressure gradient with the consideration of radial equilibrium effect to split a single shock foot into multiple weaker shock wave configuration, is investigated on the NASA Rotor 37 for promoting aerodynamic performance of a transonic compressor rotor. Numerical investigation on baseline blade and improved one with blade modification on suction side has been conducted employing the Reynolds-averaged Navier–Stokes method to reveal flow physics of ramp structure. The results indicate that the passage shock foot of baseline is replaced with a family of compression waves and a weaker shock foot generating moderate adverse pressure gradient on ramp profile, which is beneficial for mitigating the shock foot and shrinking flow separation region as well. In addition, the radial secondary flow of low-momentum fluids within boundary layer is decreased significantly in the region of passage shock-wave/boundary-layer interaction on blade suction side, which mitigates the mass flow and mixing intensity of tip leakage flow. With the reduction of flow separation loss induced by passage shock, the adiabatic efficiency and total pressure ratio of improved rotor are superior to baseline model. This study herein implies a potential application of ramp profile in design method of transonic and supersonic compressors.

Aerodynamic and structural performances of a single-stage transonic axial compressor with blade fillet radius

PurposeThis study’s investigation aims to clarify the effect of an additional geometry, i.e. a fillet radius, to the blades of a single-stage transonic axial compressor, NASA Stage 37, on its aerodynamic and structural performances.Design/methodology/approachApplying the commercial simulation software and the one-way fluid–structure interaction (FSI) approach, this study first evaluated the simulation results with the experimental data for the aerodynamic performances. Second, this paper compared the structural performances between the models with and without fillets.FindingsThis research analyses the aerodynamic results (i.e. total pressure ratio, adiabatic efficiency, stall margin) and the structural outcomes (i.e. equivalent von Mises stress, total deformation) of the single-stage transonic axial compressor NASA Stage 37.Originality/valueThis paper mentions the influence of blade fillets (i.e. both rotor hub fillet and stator shroud fillet) on the compressor performances (i.e. the aerodynamic and structural performances).

Impact of inlet flow angle variation on the performance of a transonic compressor blade using NIPC

The impact of inlet flow variations on turbomachines is common in the real world, resulting in undesirable performance deteriorations in engineering. The origin of performance changes and how to evaluate the changes require to be investigated. On use of a polynomial chaos based surrogate model method and Monte Carlo simulation (MCS), performance impact considering the stochastic variations of the inlet flow angle for a transonic compressor rotor blade is studied, and the impact mechanisms are investigated by MCS-based statistical flow analysis. A detailed grid-independent study assisted by Richardson extrapolation is first presented to illustrate the reliability of numerical solutions. Then, the implementations of non-intrusive polynomial chaos employing an adaptive sparse grid are introduced. Uncertainty quantifications of adiabatic efficiency and mass flow rate changes of NASA Rotor 67 are achieved, and the effects of different distributions of inlet flow angle variation (IFAV) and different ranges of IFAV on the performance changes under different operation conditions are investigated. Finally, statistical analysis of flow solutions is performed by MCS. The spanwise distributions of the changes in adiabatic efficiency and mass flow rate at the outlet are presented. Moreover, the flow solutions on the blade-to-blade stream surfaces at different spans are statistically analyzed, and discussions are given to demonstrate the impact mechanisms of IFAV on performance changes.

Investigation of Suction Flow Control on Centrifugal Compressor with Vaned Diffuser

In this study, the effect of suction flow control on a centrifugal compressor at operation and stall flow rates was investigated using computational fluid dynamics (CFD). The compressor geometry was reconstructed from available open source profile data and the CFD analyses have been performed on this geometry using the appropriate mesh. To validate the CFD results, the compressor performance line was acquired and compared with the experimental results obtained at the design rotational speed. Then, suction flow control was employed at various suction slot positions with different suction flow rates to improve the performance of the compressor at operation and stall flow rates. As a result of the suction flow control trials, 0.85% increase in pressure ratio and 0.8% increase in adiabatic efficiency were achieved while the compressor was running at operation flow rate. The performance improvements corresponding to the stall flow rate of the compressor were 2.5% increase in pressure ratio and 2% increase in adiabatic efficiency.

А. А. Мерти, С. Норрис, А. Субианторо. Влияние масла на мощность ротационного детандера с четырьмя пересекающимися пластинами = Murthy A. A., Norris S., Subiantoro A. Effects of Lubricating oil on the performance of a Four-Intersecting-Vane Rotary Expander / пер. с англ. М. А. Федоровой

Lubrication plays an important aspect in machinery. It seals internal leakage gaps, reduces friction and cools components. The present investigation is intended to test the effects of lubrication to an expander’s performance. Polyester lubricating oil grades (Grade 48 and Grade 68) were tested at various operating conditions. The expander prototype is of the rotary mechanism with four vanes comprising two intersecting bars. An experimental test rig was designed and built. The pressure diagram as function of rotational angle was recorded by six pressure transducers arranged in the working chamber, based on which the features of the expansion process of the four-intersecting-vane rotary expander were analysed. To start with, the working fluid was compressed air, but the eventual intended application is in a vapour compression refrigeration system. The experimental study revealed that the rotary vane expander worked steadily at different operating conditions. The experimental results showed that the expander flow rate was generally lower with a thinner oil (POE grade 46), whereas the volumetric efficiency increased from 23,2 % to 27,2 %. The adiabatic efficiency decreased from 22,5 % to 18,1 % with a higher viscosity lubricating oil (POE grade 68).