Band 40, Heft s1

To ensure that the aerothermodynamic cycle design of a turbofan engine is more accurate, efficient, and provide a reliable decision-making basis for engine designers, the multi-objective particle swarm optimization (MOPSO) method was used to optimize the aerothermodynamic performance parameters of the turbofan engine at multiple design points (MDPs). Fuel consumption rate and the specific thrust were considered as optimization targets. The thrust requirements and cycle parameter constraints under each working state were comprehensively considered to obtain the optimal performance boundary of the engine, the corresponding cycle parameters, and the correlations between different requirements and constraints. The results showed that the MOPSO algorithm could accurately and completely obtain the optimal performance boundary surface of the engine in the feasible region and the corresponding cycle parameter value. The feasible region obtained by the aerothermodynamic cycle design at MDPs was more accurate and effective than the design at a single design point.

To address the limitations of conventional groove designs in groove flow control technique, this paper optimizes the groove flow control technique for an axial-flow pump combining the design of experiment (DOE), response surface methodology (RSM), and particle swarm optimization (PSO). The sample space is designed using a combination method (OD-LHS) of orthogonal design (OD) and Latin hypercube sampling (LHS). Performance prediction models for the axial-flow pump are established using RSM. Taking the multi-condition comprehensive evaluation function as the final optimization objective, PSO is used to find the optimum groove parameters. The results show that the proposed method is effective in solving multi-condition optimization problems for grooves in axial-flow pumps. The optimal groove length, depth, and distance from the center of the impeller are 0.8, 0.05, and 0.2 times the impeller diameter, respectively, and the number is three times the number of blades. In addition, the optimal grooves effectively improve the hydraulic performance of the axial-flow pump under stall conditions. This study sheds light on the design optimization of the groove flow control technique for axial-flow pumps and other types of hydraulic machinery.

In order to study the influence of different nozzle configurations on the gas–liquid two-phase pulse detonation engine (PDE) propulsion performance, the measurement system based on the tunable diode laser absorption spectroscopy (TDLAS) technology is built to measure the velocity and the temperature, while the high frequency dynamic pressure sensor is used to measure the nozzle gas pressure. Based on the momentum principle, the contribution mechanism of unsteady gas jet on thrust is obtained indirectly by TDLAS data. The results show that the impulses of PDE with non-nozzle, convergent nozzle, divergent nozzle and convergent–divergent nozzle are 1.95, 2.08, 1.85 and 2.16 N∙s within 20 ms of the exhaust period, respectively. The analysis reveals that the impulses of PDE with convergent and convergent–divergent nozzles are larger than that with non-nozzle, while the impulse of PDE with divergent nozzle is smaller than that with non-nozzle. The research results in this paper can provide reference for the design of nozzles for PDE.

A detailed numerical analysis on a scramjet combustor is carried out by introducing an innovative shaped strut in place of the conventional strut. The design of newly added strut aids in generating intense vorticity which helps in efficient mixing of fuel and oxidizer. The air from the isolator enters the combustor at Mach 2.0, whereas fuel enters from the trailing edge of the strut sonically. In this study the flow dynamics with finite volume approach on commercial software Ansys-Fluent 20.0 to solve the two-dimensional Reynolds average Navier Stokes equation (RANS) with compressible fluid flow by considering the density-based solver with SST k-ε turbulent model. The species transport model with volumetric reaction and finite rate/eddy dissipation turbulence chemistry interaction is adopted to study the combustion phenomena and validated with the experimental results, and it is found that the interaction of the shear shock layer enhances the mixing rate by intensifying turbulence. The modified strut injector’s mixing efficiency is compared to the base strut and observed that with a 40% reduction in length, the modified strut injection technique exhibited a mixing efficiency of >95%. The combustion efficiency is then estimated streamwise, and the plot follows the same pattern as the mixing efficiency with fuel burns down completely when x  = 150 mm for the modified strut whereas x  = 200 mm for the base strut. This can compact the combustion chamber and increases the thrust-to-weight ratio. So, the innovative strut adopted can improvise the combustion efficiency.

During gas turbine engine testing, steady-state gas-path stagnation temperatures and pressures are measured in order to calculate the adiabatic efficiencies of the major turbomachinery components. These measurements are carried out using fixed intrusive probes, which are installed at the inlet and outlet of each component. The overall uncertainty in calculated component efficiency depends on the accuracy of discrete point pressure and temperature measurement. High accuracy in measurement and prediction of measurement errors has become increasingly important if small gains in component performance needs to be achieved. The recent trend is to predict component efficiencies within ±1–2%. The present work covers different Kiel designs that have been developed in a response to this demand based on a MATLAB code and experimental evaluation. A parametric study has been carried out by varying the two most critical parameters viz. Ae/Ab ratio and L/D ratio to optimize the Kiel design. These design changes will allow measurements to be made with minimum possible errors and efficiencies to be calculated more accurately over a wider range of conditions inside a low bypass turbofan gas turbine engine.

Turbojet applications benefit from accurate performance models. The aim of this study is to explore the applicability of data-mining algorithms to determine relationships between the generated thrust, the environmental conditions (free stream air-speed, inlet temperature and pressure) and the operating conditions (input fuel flow and shaft angular speed). For this purpose, experimental tests were carried out within wind-tunnel facilities using an experimental single-spool turbojet test bench. It is well-known that a large set of data-mining approaches relies on establishing linear correlations among input and output variables. The scope of this article is to assess the applicability of correlational data-mining approaches by i) an exploratory data analysis to find underlying data patterns and ii) principal component regressions to obtain a suitable predictive model for the generated thrust. Validation experiments demonstrated that the data-based model allows capturing the effects of the environmental and operating conditions with good accuracy (Root Mean Squared Error RMSE = 3.5100%), while maintaining a low complexity in the resulting structure. These results show that it is possible to generate turbojet experimental performance maps through data-mining algorithms with a correlational approach.

The present study performed a three-dimensional numerical analysis on an adiabatic flat plate with forward injection holes for multi-zone film cooling. The cooling holes were divided into three-zone, and the cold air was supplied from cylindrical holes at a velocity ratio of 0.5 and 1.5 with 30° inclination to the primary flow. The effect of multi-zone arrangement in film cooling effectiveness is studied, and a comparison between two-zone and three-zone arrangement has been made. Results show that the three-zone arrangement helps achieve better film cooling effectiveness than the two-zone arrangement due to the uniform flow of coolant at a higher velocity ratio. It also reduces the mass flow rate of secondary flow by decreasing the number of cylindrical holes in the perforated plate.

To obtain reliable and accurate experimental data in cascade testing, the influencing factors and the improving method of the flow quality of a highly-loaded compressor cascade under high incidence were investigated through a series of numerical simulations and experiments. The numerical method was validated by experimental data and agreed well at both incidence angles of 0° and 6°. Under the original upper end wall, both experimental and numerical results indicated an unsatisfactory flow quality of the cascade with an obvious nonuniformity of inlet Mach number, and the incidence of the central blade is 3.6° larger than the theoretical value. Using a small curved upper wall can reduce the severe flow separation on the upper wall and achieve a maximum improvement in flow quality under the critical installation angle, where the incidence deviation of the central blade was reduced to 2.1°. Using the combination of adjustable tailboards and a small curved upper end wall can further improve the cascade flow quality. Under the optimal angle of the tailboards, both the inflow uniformity and the outflow periodicity of the three middle blade passages the test requirements, and the incidence deviation of the central blade is reduced to 0.2°.

Compared with wall emission, gas thermal radiation is much more complicated because of its nongray and volumetric property. In this paper, a numerical method is established to calculate 3D radiative heat transfer in combustor by modelling radiative transfer as well as nongray radiative properties of combustion gases. Energy exchanges caused by thermal radiation and conduction are calculated and compared in a rectangular combustor, which shows the significant role of thermal radiation in heating fuel-air mixtures and prompting internal combustion reactions. Besides, radiative heat flux on the wall is also quite obvious although a non-contacting flow case, revealing the special challenges for thermal protections brought by radiant energy. Lastly, increasing the working pressure means much more participating species in radiative transfer process and the radiative effects will be also magnified. The numerical method in this paper provides a direct technique to analyze the role of thermal radiation in complex thermochemical reactions while the application case proves the necessity of coupling a high-accuracy radiation model when simulating combustion and flame propagation.

A creep-fatigue life prediction model based on a novel creep damage evaluation method (NCDEM) considering the multi-factor coupling effect is presented in this paper. Further, to verify the validity and practicability, the creep-fatigue life of GH4169 at 650 °C is calculated to compare with the experimental results. Ultimately, the prediction results are respectively compared with those of the creep-fatigue life prediction models based on the time fraction method (TFM), ductility exhaustion method (DEM), and strain energy density exhaustion method (SEDEM). Consequently, the prediction results are distributed in ±1.5 times dispersion band, which elucidates the creep-fatigue life prediction model proposed based on the NCDEM has the best ability.

Engine transient control has been challenging due to its stringent requirements from both performance and safety. Many methodologies have been proposed such as conventional schedule-based methods, linear parameter varying, multiobjective optimization and evolutionary computations etc. These approaches have been well-established and led to a series of significant results. However, they are either not providing limit protection or requiring exhaustive computational resources, particularly when generating results into full flight envelope applications. Consequently a compromise between limit protection and computational complexity is necessitated. This note considers a sequential quadratic programming (SQP)-based method for full flight envelope investigations. The proposed method can provide important design guidance and the corresponding claims are validated through detailed analysis and simulations.

The present work reports the influence of the 1.5-stage turbine flow field by the front and aft rim seal flow. The interaction between the front and aft purge flow and the mainstream of a 1.5-stage turbine was numerically simulated, and the influence of the front and aft purge flow on the downstream vane was analyzed separately. The results show that the front purge flow is distributed at the higher radius of second vane inlet, which changes the position of the blade hub secondary flows, and the aft purge flow is distributed at the low radius. The purge flow at different locations in the aft cavity exit forms shear induced vortex, pressure and suction side legs of the egress, which converges with the suction and pressure side legs of the horse vortex to form vane hub passage vortex. The increased purge flow rate in both the front and aft cavities significantly increases the sealing effectiveness of the rim seal, but also causes a reduction in turbine efficiency. The combined effect of the front and aft purge flow reduces the turbine efficiency of the end-wall structure by 0.3619, 0.9062, 1.5004, 2.0188 and 2.509% at IR  = 0, IR  = 0.5%, IR  = 0.9%, IR  = 1.3% and IR  = 1.7%.

Bladed disks are important components of gas turbine engine. Rotor disk spool drum assemblies of gas turbine engine constitute 20–25% of total engine weight. Increasing thrust-to-weight ratio and engine life is paramount for designers. Blisk reduces significantly weight of rotor, compared against conventional disks for aero engines. This paper brings out specific challenges faced while re-designing bladed disk into blisks including structural integrity aspects under various operating loads. This paper presents a case study on re-design of typical compressor bladed disk into a blisk, without changing the flow path or airfoil configuration, within space constraints. Weight reduction of rotor disk is carried out using shape optimization technique. Blisk configuration is derived from existing bladed disk general arrangement. This paper describes methodology of weight optimization of blisk using ‘HyperStudy’ tool considering static and dynamic 3D models with ANSYS solver. APDL fatigue life macro is developed for fatigue life prediction, using strain-life approach. In this paper 3D bladed disk, baseline and optimized 3D blisk modal analyses results are used to ensure minimum interferences for engine operating conditions. The developed methodology for optimization can be appreciated by significant weight reduction (30%), while meeting design criteria and increased fatigue life.

Flutter, an aeroelastic blade vibration phenomena, experienced by the fan of an developmental aero gas turbine engine, result in blade failure. Hence, suitable flutter detection instrumentation is required during engine testing. Flutter signature capture from revolving blades is a challenging task that necessitates either a complicated strain gauge-based rotating instrumentation or a noncontact tip timing system. Authors investigated a unique way for identifying, measuring, and validating flutter signature by assessing wall static pressure pulsations produced during blade tip transit across a casing mounted high bandwidth sensor during this research. The authors devised a mathematical model to explain signal spectrum components that feature both amplitude and angle modulation properties at the same time. The theory was tested using first-stage fan rotor blades that were fluttering in the first flexural mode (1F) and forming the second nodal diameter (2ND). The approach’s estimated blade deflection was compared to measurements taken using a traditional tip timing method up to 7 mm and determined to be within 1% inaccuracy. This research provides a low-cost, easy alternative technique for measuring flutter during engine development testing.

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.

The field of Prognostics and Health Management (PHM) in industries is gaining greater popularity to achieve high reliability by shifting the preventive maintenance to predictive maintenance. Estimation of Remaining Useful Life (RUL) is an effective prognostic measure that forecasts the health state of machine based on degradation modelling and condition monitoring. This article proposes a novel and robust methodology that uses Reduced Affinity Propagation (RAP) clustering technique that extracts representatives from the temporal signals measured through various heterogeneous sensors to predict the RUL using Echo State Network (ESN) with dynamic lateral inhibiting connections. The main advantage of the proposed model is that it does not overlook the features from the degradation signals and also learns the natural mapping among the representative points from the integrated sensor value. This approach is verified using CMAPPS dataset to show hopeful results in predicting the RUL of aircraft turbo fan engine. Also, this methodology can be a deployed as a tool in real time industrial applications to schedule predictive maintenance activities.

To solve the problems of the nonlinear damping force in the traditional squeeze film damper (SFD), a novel structure of G-type integral squeeze film damper (GISFD) based on ISFD is proposed for the first time. The finite element model and test rig of the ball bearing-rotor system are established to explore the influence of GISFD and ISFD on the dynamic characteristics of the unbalanced rotor system. The results show that both GISFD and ISFD can change the critical speed of the rotor system, reduce the bending strain energy of the shaft, and reduce the bearing dynamic load of the rotor system. Through comparison, it is found that the effect of GISFD is more obvious. The experimental results show that, compared with the unbalanced rotor system without damper, the peak-peak value of amplitude in the rotor system with GISFD and ISFD at 3000 rpm is reduced by 25.53 and 15.81%. The amplitude in the disk at the first-order critical speed is effectively reduced, and the reduction range reach 52.01 and 35.44%, respectively. GISFD has a more significant effect of suppressing unbalanced vibration, and has superior vibration damping performance when compared with ISFD.

The use of micro gas turbines (MGTs) for once-off and unique applications means that a rapid turnaround design process is required. The development of a MATLAB ® application-based program for the initial design and performance analysis of radial and mixed flow compressors is discussed. The program code is based on one dimensional (1D) mean line flow and loss model theory. For verification, 18 test compressors were developed, covering a wide range of design velocities, mass flow rates and meridional exit angles. Predicted performance results were verified using Numeca/FINE™ Turbo software. Initial comparisons between 1D and Computational Fluid Dynamics (CFD) results did not match well. The 1D software over-predicted compressor performance and provided poor choke prediction. Consequently, empirical models to correct these deviations were derived and implemented into the 1D Application code. The updated 1D code provided a mean choke prediction difference of 1.59% compared to an initial 14.98% difference at the design point. Mean total-to-total isentropic efficiency and pressure ratio reduced to differences of 0.74 and 1.24% from values of 11.23 and 9.31% respectively.

This paper focuses on the investigation of the blade deformation and tip clearance effects of a high-flow-coefficient centrifugal compressor. Blade deformation caused by the pressure load and centrifugal load under operating condition is considered by fluid/solid interaction method to simulate the stage performance accurately. Hence the numerical results are in good agreement with the experimental data. The blade deformation is then analyzed in detail. Also, the effect of the size of tip clearance is studied by evaluating the performance with five various gaps. The result shows that the increment of the tip clearance from 0 to 2.0 mm leads to a 5.14% peak efficiency decline. However, increasing tip gap contributes to enlarge the stable operating range in terms of stall margin and choke margin.

Turboprop engines require an exhaust nozzle or stub to duct the engine exhaust flue gas outboard of the aircraft. The design of these exhaust stubs are dictated primarily by the aircraft’s configuration. In pusher aircraft, the exhaust stubs are designed to minimize the exposure of the flue gases from the engine exhaust on the propeller blades and fuselage. A fluid-thermal-structure coupling analysis is performed to understand the thermal effects of the engine exhaust jet flow on the thermo-mechanical behavior of pusher configured light transport aircraft propeller and structure. The steady thermal flow field of the aircraft with forward and reverse thrust, in which propeller blade angle variations were analyzed for different aircraft speed. The present work investigates a three-dimensional analysis of flow around the nacelle-airframe and the effect of exhaust flue gas impingement on the propeller blade surface. Based on the insights from the numerical results, the designed exhaust duct was integrated on the aircraft and carried out ground static and flight testing for various flight operating conditions in which propeller blade and fuselage surface temperature were measured. Numerical and experimental results are compared and validated for certain flight conditions and found satisfactory.

To further reveal the interaction between crossflow and jet, single oblique jet impingement in crossflow is studied. The influences of Reynolds number (Re = 3750–15,000), oblique angle ( θ  = 50°–90°), jet intake mode (jet normal to inlet plane and jet along hole axis), crossflow ratio (CR   = 0.1–0.3) and temperature ratio (TR   = 1.0–1.2) are considered. For both jet intake modes, with the decrease of  θ , the peak value of Nusselt number increases and its location shifts upstream. The Nusselt number for case of jet normal to inlet plane is higher than that of jet along hole axis with wider lateral coverage range. The CR affects the convection heat transfer between crossflow and target surface, and it also affects the interaction between crossflow and jet. Synthetically, the best heat transfer appears at CR = 0.2. The Nusselt number decreases sharply when TR goes from 1.0 to 1.1, while the decline trend slows down when TR goes from 1.1 to 1.2.

With the purpose of investigating the effect mechanism of injection on flow separation of the S-duct, different single-hole schemes were investigated and compared with double-hole schemes. Results show that, the performance of S-duct can be improved by using injection. The optimal scheme in this study is a double-hole injection scheme with two holes located at the same axial position. Flow separation reduction and a 16.9% reduction of loss coefficient were achieved by injection with an injection coefficient of 0.46% in each hole. The flow mechanisms are that, firstly, high momentum fluid is injected to separated flow by air injection; secondary, high momentum flow is transported to flow near downside wall by injection vortex. The position effect, injection flow rate effect and hole shape effect were also discussed. For double-hole scheme, the scheme with two holes located different axial positions generates a stronger vortex by mixing two injection vortexes and enhances secondary flow. Though the flow separation is reduced, a severe nonuniform flow field at outlet is formed. Due to the less swirling flow achieved at outlet of S-duct by Double-y scheme, it can offer a more uniform flow field for downstream compressor and has a better control effectiveness.

The paper outlines a finite volumes refined mathematical model of the working gas flow in the flow path of the three stage modern single shaft gas turbine engine that can be used in floating power plants. Such mathematical model based on the finite volumes of hexagonal-type was constructed using the three-dimensional Navier–Stokes equations for the case of viscous working fluid flow. For the problem solution such boundary conditions as “inlet’, “outlet” and “wall” have been used. The calculation is carried out in a non-stationary setting with a time step of 1.5974 × 10 −6  s, which corresponds to the angle of rotation of the rotor, relative to the stator, of 0.09°. The total number of time iterations is 350. Also, it was shown that the variation field of pressure on the blades feather surfaces and the gas flow velocity due to rotation are the critical factors, causing the blades vibration. The result was confirmed with the experiment. The obtained results would be used as a base for further investigations of gas flow pressure field on the blades surface, because the gas flow pressure are key factors, causing the rotor forced vibration, and as initial data for their fatigue strength and crack study.

Aero engine performance deterioration highly influences its reliability, availability and life cycle. Predictive maintenance is therefore a key figure within Industry 4.0, which guarantees high availability and reduced downtime thus reduced operational costs for both military and civil engines. This leads to maintenance on demand and needs an effective engine health monitoring system. This paper overviews the work carried out on aero engine health monitoring, diagnostic and prognostic techniques based on gas path performance parameters. The inception of performance monitoring and its evolution over time, techniques used to establish a high-quality data base using engine model performance adaptation, and effects of computationally intelligent techniques on promoting the implementation of engine fault diagnosis are reviewed. Generating dependable information about the health condition of the engine is therefore a requisite for a successful implementation of condition-based maintenance. Based on this study, further research can be attempted to predict residual life of critical components using degradation pattern from aero engine performance data bank which will be an invaluable asset for engine designers as well as for operators.

Leading edge (LE) plays a prominent role in compressor flow. The asymmetric leading edge (ASYLE) has shown superiorities to symmetric LE in blade aerodynamic performance. However, the influencing rules of ASYLE design parameters are still ambiguous. In this work, numerical calculations were conducted to investigate the influencing effects of LE point curvature and position. The results show that the operating range of ASYLE blades expand with the decrease of LE point curvature, which helps to moderate LE flow acceleration, while the LE point position mainly affects the operating range. It is also revealed that the SSLE curvature peak is supposed to be close to LE point, and the maximum value of PSLE curvature should be restricted.

The changes of operating conditions can lead to variations in the aerodynamic conditions of the turbine blades. Then numerical studies are conducted to study the aerothermal performance of the turbine blade with a squealer tip at various Reynolds numbers ( Re ) and exit Mach numbers ( Ma ) with moving endwall. Besides, the effect of the rim height on the squealer tip is studied. The increasing of Reynolds numbers at Mach number = 0.78 enhances the heat-transfer at the tip of blade entirely, however, it has little impact on the leakage flow of the blade tip. The total heat flux increases with Mach number increasing from 0.78 to 0.96 and changes slightly with a further increasing of Mach number at a design Reynolds numbers. The height of squealer rim affects the tip leakage flow in the cavity and tip leakage vortex, and then affects the heat transfer distribution and the heat load of the blade tip.

A variable stator vane (VSV) with nonuniform partial radial gaps is numerically investigated in an annular compressor cascade. Five adjusting angles ( α  = −5.5°, −5°, 0°, 5° and 10°) are chosen to simulate different working conditions. The VSV is adjusted in negative direction means that the stagger angle increases and the VSV is more closed. Since the VSV is installed in an annular cascade, the heights of the partial gaps are nonuniform and reaches 3.8% of the span. Results show that as the VSV is adjusted negatively, the total pressure loss rises by 20.3% and a huge area of corner separation appears. The outflow angle is also more distorted along radial direction. Comparisons with a fixed-configuration stator show that although the VSV would cause higher loss, it indeed plays an important role in adjusting the outflow angle.

Counter-rotating axial-flow compressor (CRAC) is a promising technology to enhance the thrust-to-weight ratio of aero-engines. Self-recirculating casing treatment (SRCT) is an efficient flow control technique for increasing stall margin in conventional compressors. With the purpose of investigating the applicability and mechanism of SRCT in the CRACs, a two-stage CRAC is selected to investigate the stability enhancement mechanism of SRCT and its effect on the unsteady flow near the rotor tip, and the effect of injector location on the stability improvement capacity of SRCT is also studied. Results show that about 7.73% stall margin improvement can be achieved by configuring the SRCT on the near rotor top, and the injector location also has a significant influence on the stability expansion potential of SRCT. The SRCT delays the stall occurrence by weakening the intensity of tip leakage flow (TLF) and restraining the leading-edge spillages of TLF. The SRCT reduces the unsteady interference between the adjacent rotors by receding the disturbance of the upstream wake and inhibiting the potential flow effect of the downstream. Furthermore, the SRCT reduces the self-excited oscillation frequency of TLF and damps its fluctuation amplitude.

Aiming at the problem of frictional heat generation under high pressure and high speed of the piston ring, the V-groove piston ring is proposed. The finite element analysis method is used to compare and analyze the temperature distribution law of the traditional piston ring and V-groove piston ring. The experimental study on the temperature rise of the piston ring seal is carried out by using the high-pressure and high-speed rotating sealing experimental bench. The experimental results are consistent with the simulation results. The results show that when the rotor speed increases by 1000 rpm, the temperature rise of the traditional piston ring increases by 18.4%, the V-groove piston ring increases by 17.2%. For every 0.5 MPa increase in the pressure applied to the piston ring, the temperature rise of the traditional piston ring increases by 9.2%, the V-groove piston ring increases by 5.8%. The same under the working conditions, the temperature rise of the V-groove piston ring is reduced by 21–29% compared with the traditional piston ring, and the leakage of the V-groove piston ring is reduced by 12–19% compared with the traditional piston ring. The synergistic design of low temperature rises and low leakage of the piston ring.

The accumulation of catalyst dust at the honeycomb seal for a long time will lead to the honeycomb seal cell blockage, weaken the seal effect of the honeycomb seal, and directly affect the safe and stable operation of the flue gas turbine. Therefore, a research on the influence of honeycomb cell blockage on the seal leakage characteristics is carried out. The flow field model of honeycomb seal is established, and the influence of cell blockage on the leakage characteristics of honeycomb seal is analyzed. A seal leakage characteristic experiment bench is built, and the numerical simulation results are verified by the experiment bench. Compared with the honeycomb seal without blockage, the leakage of the honeycomb seal with a blockage rate of 25% increases by about 6.5%, while the leakage of the honeycomb seal with a blockage rate of 25% obtained by the experiment increases by about 6.3%. The numerical simulation results are in good agreement with the experimental results. The results of this research provide theoretical support for revealing the seal leakage characteristics of honeycomb seals under blockage faults.

As an essential component, the bleed system plays a critical role in supplying turbine cooling air, guaranteeing stage matching, pressurizing the cabin, and de-icing at the wing and engine inlet. However, the extraction of the bleeding air from the compressor causes the engine efficiency degradation and thrust deficit. Therefore, flow control based on bleed is conducted to compensate the bleed induced disadvantages. The influence of the circumferential bleeding slot location on the tip leakage vortex and passage vortex controlling in a compressor cascade with the tip clearance is numerically studied using large eddy simulation. Three bleed configurations and the smooth casing configuration are investigated. 17.11% loss reduction is obtained through bleeding at 10% c x upstream of the blade leading edge with a bleeding rate of 2.76%. The vortex structures and flow patterns are compared and analyzed to reveal the controlling mechanism. Subsequently, the axial vorticity and loss evolution is discussed, and the interaction between the primary flow and bleeding air is revealed. It’s found that bleeding slot placed within the blade passage is exposed into a highly static pressure gradient, and this causes the bleeding air flows into and spills out the bleeding slot and leads to unnecessary loss. Moreover, the influence of large bleeding rate and inlet boundary layer is assessed.

To improve the prediction accuracy of profile loss at low Reynolds number, a typical low-pressure turbine cascade T106D-EIZ was selected to numerically investigate the effect of Reynolds number on turbine cascade flow. A detailed analysis of profile loss was performed and a profile loss model considering the low-Re effect was developed. Results showed that the incidence angle has a great effect on the inlet and outlet Mach number at low Reynolds number, and the variation of inlet and outlet Mach number further affects the blade profile loss. A correction factor was introduced to consider the effect of incidence angle and Mach number on the profile loss. The profile loss coefficient and stalling incidence angle were both extended to lower Reynolds number based on the numerical results. A Smart Through Flow Analysis Program (STFAP) was developed using the finite volume method to solve the circumferentially averaged Euler equations of S2 surface. Aerodynamic performance of E3 5-stage low-pressure turbine was predicted by STFAP coupled with low-Re profile loss model. Compared with K-O model, the prediction accuracy of efficiency of low-pressure turbine last stage is improved by nearly 1.1 percentage points when the 5-stage low-pressure turbine is in a low Reynolds number state.

Aiming to optimize the mode transition control schedule of the adaptive cycle engine (ACE), an optimization method based on a gradient algorithm was proposed. During the mode transition, the compressor surge margin, the total turbine inlet temperature and the combustion chamber fuel-air ratio are not exceeded. The integration of thrust by time was selected as the optimization objective. The geometric parameters were updated via the optimization target derivative. The speed limit of the geometric adjustment mechanism was also considered in the optimization process. The control schedules of two mode transition processes, from triple to double bypass mode and from double to triple-bypass mode were optimized. The optimization method presented in this paper can be applied to various mode transitions and the geometric adjustment speed meets all the constraints.

This paper reports the numerical investigations to analyze the effect of the application of Gurney Flaps on various configurations in a Low-Pressure Turbine (LPT) cascade in mitigating the laminar flow separation during low Reynolds number operations. T106 LPT blade of chord 60 mm has been selected for the present study. Gurney Flaps of flat type and quarter round type of a particular height of the chord length are selected. Flaps are provided near the blades’ trailing edges for numerical analysis. The numerical computations are performed using STARCCM+ software, and the K-ω SST turbulence Model is used for turbulence closure. The studies are performed at various Reynolds numbers ranging from 37,500 to 138,750 to understand the application of various configurations of Gurney Flap vis a vis turbine blade without Gurney Flap. Performance parameters such as lift coefficient, drag coefficient, aerodynamic efficiency, and static pressure distributions over the blade surface are used to analyse the alterations in the performance of aerodynamic characteristics of the LPT Blade. All the Gurney Flap configurations improved the overall blade aerodynamics. Among these configurations, the flat Gurney Flap configuration is found to be superior as it improves the aerodynamic efficiency of the blade.

Aiming at the problems of large leakage and high friction heat generation of the piston ring under high pressure and high speed, based on the previous research on the V-shaped groove piston ring, the research on the design method of the V-shaped groove piston ring is carried out in depth. The numerical analysis is carried out focusing on the number of V-shaped grooves and the depth of the V-shaped grooves of the piston ring. The influence law is verified and analyzed by using a high-pressure and high-speed rotating test bench. The simulation and experimental results show that the leakage of 4 grooves and 8 grooves piston ring is reduced by 14.9 and 28.5%, and the temperature is reduced by 10.9 and 4.7%; When there are 8 V-shaped grooves on the piston ring, the leakage is reduced by 28.5 and 21.7%; when the V-shaped groove depth is 0.04 and 0.08 mm, the temperature is reduced by 4.1%. However, when the V-shaped groove depth is 0.12 mm, the leakage is increased by 2.7% compared with the traditional piston ring. Appropriate groove number and groove depth can realize the coordinated design of low leakage and low friction temperature rise.

The passive control of jets using vanes as vortex generators is studied by numerical simulation in this paper. The vanes are positioned inside the nozzle near the exit, inclined to the flow with the longitudinal direction of the jet. Two configurations namely, straight vanes ( k  = 0 mm −1 ) and curved vanes ( k  = 0.05 mm −1 ) are considered. Curvature k is defined as the reciprocal of the radius of the vanes. The blockage due to the presence of the vanes is 0.5%. The total pressure variation along the jet centreline and along the radial distance is determined from nozzle exit at a Mach number of 0.4, 0.6 and 0.8. It is found that the vanes cause faster decay of the jet, both in the near field and far field compared to the base nozzle (plain circular nozzle) and the curved vanes perform better than the straight vanes in promoting the jet mixing. A maximum of 54% reduction in jet potential core length is achieved by the curved vanes and the jet becomes asymmetrical due to the presence of the vanes inside the nozzle, as observed in the radial pressure decay plots and Mach number contours.

The variations of inlet environment parameters can make significant effects on the compressor performance. This paper aims to investigate the effects of inlet total pressure and total temperature changes on the rated condition performance of a nine-stage HPC. Different cases of total pressure and total temperature boundary conditions at this compressor inlet are studied by 3-D numerical simulations with experimental validations. The numerical results confirm that the variations of inlet total pressure and total temperature make different effects on the rated condition performance of compressor. The overall performance parameters, such as the corrected mass flow and isentropic efficiency, will increase with inlet total pressure increasing and decrease with inlet total temperature increasing by different change rules. The flow similarity is also investigated by comparing the calculated results of critical quantities in different cases. The results indicate that the rising inlet total pressure can increase the Reynolds number and it is beneficial to reduce the viscous influence so that it is available to improve the performance; the rising inlet total temperature can decrease both the specific heat ratio and Reynolds number so that it will lead to the compressor performance decline inevitably.

In view of the problem that the vibration of the centrifugal impeller of the compressor exceeds the standard when a micro turbojet engine crosses the first-order critical speed, this research proposes a vibration control technology for the engine rotor system based on the new integral squeeze film damper (NISFD) according to the unique structure and assembly characteristics of the rotor bearing seat of the micro turbojet engine. Through numerical analysis and experimental research, the dynamic characteristics of NISFD are analyzed, and the vibration reduction mechanism of NISFD is revealed. The engine test verification based on NISFD is carried out. The results show that NISFD has good damping characteristics, which can improve the bending stress and strain energy distribution of the rotor. The engine test proves that NISFD can well solve the problem of excessive vibration of the engine when passing through the first-order critical speed, greatly reduce the vibration level within the working speed range, and significantly improve the reliability of the engine during operation. It provides a solid theoretical basis for the design of aeroengine, and provides a reference for the practical engineering application of NISFD in aeroengine.

Aiming at the problem of aerodynamic coupling interference between disturbance of the built-in measurement probe and internal flow of high load compressor, the experiment of axial position and structure dimension of cylindrical probe support effect on the performance of the compressor cascade was carried out on the transonic plane cascade test facility. The effect characteristics of the distribution of probe support on aerodynamic performance of downstream cascade under different inlet Mach numbers were analyzed through the experimental data. The experimental results show that the axial position of probe support has little impact on the total pressure loss coefficient of cascade under the low-speed inlet flow condition. However, the variety of the axial position of probe support has a great impact on the total pressure loss coefficient of cascade under the high-speed inlet flow condition. With the decrease of the axial distance between probe support and cascade, the total pressure loss coefficient of cascade increases. Under different inlet flow conditions, the variety of the structure dimension of probe support has a significant impact on the total pressure loss coefficient of cascade. With the increase of the structure dimension of probe support, the total pressure loss coefficient of cascade increases.

The characteristics of a sonic under-expanded coaxial jet with lip thickness 1.5 D p (where D p is the exit diameter of primary jet equals 10 mm) with the primary jet operating at nozzle pressure ratio (NPR) of 3, 4 and 5. For NPR 3 operating primary jet, the secondary jet operates at NPR 2.5, 1.4 and 1.27. For primary jet NPR 4, the secondary jet operating NPR is 3.2, 1.6 and 1.4. For NPR 5 primary jet, the secondary jet NPR is 3.8, 1.89 and 1.52. The study is performed using a co-flow nozzle of bypass ratio (BR) 6.4, 1.4 and 0.7. The core length of the primary jet is used as a measure to quantify the mixing of the primary jet in the presence of coaxial jet. The shock structure present in the near field was viewed using shadowgraph technique. Centreline pitot pressure distribution, radial spread and waves present in the jet core were analyzed. The results show that the mixing associated with the high bypass coaxial jet is superior to the low bypass coaxial jet. This mixing superiority associated with high bypass coaxial jet prevails all levels of expansion.

Flutter-induced fatigue failure investigation of the fan blades of aero-engines necessitates extensive testing. During engine ground testing, strain gauges on rotor fan blades and casing vibration sensors were employed to investigate structural dynamic aspects. The correlation between strain sensor signals and fan casing vibration signals allowed the diagnosis of fluttering fan blades. For automated flutter detection during engine development testing, a machine learning-augmented information fusion methodology was developed. The method analyses casing vibration signals by extracting time-domain statistical features, intrinsic mode function characteristics through empirical mode decomposition, and recurrence quantification features. Feature vectors obtained from a relatively large set of engine tests were subjected to dimension reduction by applying machine learning techniques to rank them. Reduced feature vector space was labelled as “flutter” or “normal” based on the correlation of rotor strain gauge signals. In addition, the labelled feature vectors were employed to train classifier models using supervised learning-based algorithms such as Support Vector Machines, Linear Discriminant Analysis, K-means Clustering, and Artificial Neural Networks. Using only vibration signals from the casing, the trained and validated classifiers were able to detect flutter in fan baldes with a 99% probability during subsequent testing.

The present study reports the numerical investigation of the compressor cascade. To minimize the separation phenomenon in the compressor cascade, a passive flow control device i.e. Micro Vortex Generator (MVG) is utilized. MVG is a very simple and lightweight attachment mounted infront of the leading edge of the cascade blade. Due to being passive in nature, it neither consumes power nor requires any external device to actuate. The numerical simulations were carried out on a highly loaded compressor cascade at an angle of incidence of −1° under subsonic conditions at Mach number 0.2. The profile of the compressor cascade blade was double circular arc (DCA), unsymmetrical and cambered at 40°. Two different types of micro vortex generators were mounted infront of the leading edge in the compressor cascade to control the secondary flows since secondary flows were responsible for various losses in cascade. To analyze the flow under incompressoible state of air ( M  = 0.2), Star CCM + software has been used. To simulate the flow under turbulent condition, k - ω SST turbulence model was used. A velocity profile of 25 mm boundary layer thickness was extracted and used as an input in the compressor cascade. Mounting of MVG on compressor cascade enhanced drag but also increased lift. Total pressure loss coefficient (TPLC) was calculated to compare the losses. The aerodynamic efficiency in terms of coefficient of lift and coefficient of drag has been used to study the effect of MVG over cascade. It is found that there is reduction in total pressure loss coefficient (TPLC) for trapezoidal and curved trapezoidal types of MVGs and the decrease in percentage are 2.17 and 8.86%, respectively. Alos, aerodynamic efficiency is increased by mounting trapezoidal and curved trapezoidal types of MVG and the increase in percentages are 2.03 and 3.10%, respectively.

A quantitative study is carried out to explore the air/mist-coupled cooling in the cooling section incorporating a spray device, aiming to conduct a comparative simulation on the effects of two radial spray arrangements (Apparatus A and Apparatus B) on the change of flow field. The results show that an even spray arrangement leads to improved flow performance owing to the large temperature drop in the cooling section. Compared with Apparatus A, Apparatus B enables a 1.60–2.25% higher temperature drop coefficient and a 1.1–3.0% higher water evaporation rate. Six high-altitude conditions are simulated to find that the spray device and water injection are the main factors causing flow loss, whereas the additional mass flow caused by water vapor reduces the amount of loss. The highest temperature drop coefficient is observed for Case 6 (27.3%), in which there is a more uniform nozzle arrangement. A slight difference in terms of total pressure drop coefficient between Apparatus A and Apparatus B is observed. Moreover, the highest total pressure drop coefficient (5.8%) at the location of the spray system is observed for Case 6, which has the largest injection rate and highest inlet air Mach number.

To solve the poor precision and low efficiency of manual grinding de-weight dynamic balance of turbopump rotor, the laser de-weight dynamic balance technology of rotating machines is adopted. Combined with the advantages of laser de-weight, such as, fast processing speed high machining precision and good machined surface quality, the smooth transition of the machined surface was realized by means of hierarchical machining. According to the characteristics of the turbopump rotor structure, the equivalent transformation model for laser de-weight turbopump rotor was built to convert the unbalance into laser processing parameters, which achieve the rapid and accurate de-weight of turbopump rotor. It focuses on solving the repeatedly loading and low accuracy of manual grinding dynamic balance. To verify the effectiveness of the proposed technique, laser de-weight experiments were carried out on two turbopump rotor test pieces. A dynamic balance test bench was set up to measure the rotor residual unbalance at 1800 r/min test speed, which is low dynamic balance. The results show that the machined surface transition was smooth, the residual unbalance was small and the machining efficiency was high. The effect of laser de-weight dynamic balance reached the expectation. It is an universal technology which is not limited to the dynamic balance of turbopump rotor, but also applicable to other rotating machinery..

Based on the flow field characteristics of the missile body, an integrated design method for missile and inlet with two side layout is proposed under asymmetric inflow. The result of numerical simulation shows that shock waves hit on cowl lip on the symmetry plane at the design condition, which verifies the method of integrated design. And the flow characteristics of inlet under the influence of the missile body are analyzed. Under the influence of asymmetric incoming flow and missile body, the first shock wave surface of the two-dimensional inlet presents the characteristics of a three-dimensional concave surface, and there is a pair of asymmetric vortex structures in the inner flow path. Finally, a bleed cavity of self-adaptively adjusting bleeding by vortex is proposed to improve the performance of inlet and broaden the inlet start and attack angle boundary. The minimum Mach number of the inlet start is reduced from 2.8 to 2.3 at ɑ  = 6° and the maximum attack angle of inlet start is widened from ɑ  = −1° to ɑ  = 3° at Ma = 2.1.

Starting from a component-level nonlinear model of a turboprop engine, the high-pressure turbine speed and power turbine speed output data at six steady-state operating points are linearized and fitted, and a turboprop engine state variable model is established. Based on these state variable models, the Proportional Integral Derivative (PID) control method, the augmented Linear Quadratic Regulator (LQR) control method and the Linear Quadratic Gaussian/Loop Transfer Recover (LQG/LTR) control method are used to design the controllers respectively, and the relative converted speed of the high-pressure turbine is selected as the scheduling parameter of the Linear Parameter Varying (LPV) model, and the controller is called to control the turboprop engine’s non-linear speed. Linear model for large envelope control. Finally, the control effects of the above three control methods are compared and analyzed, and their advantages and disadvantages are compared. The simulation results show that the LPV controller designed based on the LQG/LTR method is more effective than the controllers designed by the other two control methods on the nonlinear turboprop model.

Since Professor Greziter first proposed a theoretical model to predict the dynamic behavior of a compression system in 1976, the contribution of the volume effect to compressor flow instabilities has been widely studied, but the role of volume effect on the compressor performance during acceleration and deceleration has not received much attention. Therefore, starting from the Greitzer lumped parameter approach and integrating with real engine simulated components to improve the Greitzer’s model, an engine transient simulation model accounting for the compressor volume effect (referred to as the unsteady model) is developed in this paper. Based on a real turbofan engine, transient examination comprising acceleration and deceleration has been conducted for the validation of the unsteady model. The simulation results show better agreement with the experimental data compared with the transient simulation without considering the compressor volume effect (referred to as the quasi-steady model), which confirms the importance of introducing the compressor volume effect into the engine transient modeling. And the cause of the deviation between unsteady model and quasi-steady model are further explained by the compressor inlet and outlet mass flow curve. The results may further contribute to the development of engine transient model.

During take-off of a vertical take-off and landing (VTOL) aircraft, ground effects can cause a downward force on the aircraft body and wings. The downward force could substantially reduce the payload of the aircraft, which is undesirable. This paper investigates the ground effects related to VTOL applications with distributed propulsion. A slot jet is used to simulate the distributed propulsion system. A model of a wing with a slot jet placed near the trailing edge of the wing is investigated. The slot jet is almost perpendicular to the ground to provide a vertical thrust. Experimental and numerical methods are used to investigate the aerodynamic performance of this model. Theoretical analysis is carried out to understand the formation mechanism of the low pressure region on the lower surface of the wing, which causes the downward force. The flow physics of the jet inducing ground vortex is investigated. It is found that the convection term in the ground vortex area is the main source of the pressure reduction. Based on the flow mechanism, a redesigned configuration is proposed to reduce the negative effect of the ground vortex. The flow structures such as the tip vortex are also investigated.

This work presents an approach for sensitivity analysis of turbine cooling blade with surface thickness uncertainties, combining mesh deformation method, neural network model and multidisciplinary analysis. Normally, for even tiny shape changes, conventional geometry-based method failed easily during the auto-processing analysis. Therefore, mesh deformation method was utilized to capture the tiny size changes in the multidisciplinary analysis for both the fluid and the structure meshes. The neural network model is constructed by design of experiments to reduce the computational cost. Sensitivity analysis of the multidisciplinary system of blade is performed by numerical difference algorithm with the neural network model. Results showed that the proposed method was effective and practical in engineering.

In order to improve the effectiveness of compressor efficiency measurement results under the influence of complex factors, several research work have been carried out, such as inverse connection of thermocouple suitable for efficiency measurement under low pressure ratio and small temperature rise conditions, torque efficiency correction considering the mechanical loss of high speed gearbox under various working conditions, and identification of the impact of seal cavity leakage flow on efficiency measurement. Compared with the conventional connection method of thermocouple, the inverse connection method of thermocouple shows obvious advantages in efficiency measurement under low pressure ratio and small temperature rise conditions. At medium and low speeds, the conventional simplified correction method will lead to the high measurement results of torque efficiency after installing the high speed gearbox. The accuracy of torque efficiency measurement can be improved by adopting the improved correction method. Increasing the oil supply pressure of the high speed gearbox or reducing the oil supply temperature will increase the mechanical loss of the high speed gearbox. The air flow in the compressor bearing seal chamber will leak into the flow passage through the labyrinth clearance, resulting in discontinuous changes of the inlet/outlet flow of compressor.

Upstream vortex has a significant effect on the secondary flow structure of the downstream turbine in the stage environment. This study investigates the secondary flow structure with non-axisymmetric endwall profiling (NAEW) under the interaction of co-rotating incoming vortex (V ic ). A half-delta wing vortex generator is utilized to model V ic . The turbine cascade case which exhibited maximum reduction of the cascade loss with NAEW under no incoming vortex is studied. The mechanism of loss reduction with NAEW under the interaction of V ic is analysed. V ic could decrease the secondary flow near the endwall region by affecting the horseshoe vortex transport in the cascade. However, its loss reduction was lower than the loss increments of V ic itself. The arrival of V ic at the leading edge of the cascade increased the strength of the horseshoe vortex, resulting in a significant increase in loss. Under the interaction of V ic , NAEW decreased the blade loading near endwall region, which resulted in the reduction of cascade loss.

As an inner heat source, the propulsion system of aircraft obviously influences the rear fuselage thermal status and thus produced infrared signature. In this paper, a coupled flow-heat simulation method is established to calculate the inner and outer flow fields of a realistic aircraft exhaust system. Considering multi heat transfer modes and fluid-solid thermal interactions, the method is examined by available experimental data. Different nozzle layouts are then provided to analyse the influence of inner hot parts including combustion gas and solid walls on the outer fuselage temperature changes. Thermal radiation is individually researched and analysed since radiation flux heats the outer fuselage skin significantly, especially in the cold convection channel. The simulation results indicate that the appearance of intermediate shield device effectively suppresses the outer fuselage temperature-rise and therefore resulting infrared signatures. In comparison, the exhaust system equipped with intermediate shield could decrease outer fuselage infrared signatures by 69 and 40% for 3–5 μm band and 8–12 μm band, respectively.

A 3D numerical analysis on an adiabatic flat plate for multi-hole trench cooling with forward, backward and mixed injection holes is performed in the current investigation. The numerical setup is validated before the performances of different cooling configurations are compared. The effect of three different multi-hole trench arrangements, square-diamond, long-diamond, and super-long-diamond with constant perforated percentage (3.27%), on film cooling performance is studied at blowing ratio 1.0. The row-to-row interaction between coolant jets and mainstream is analysed, and lateral film cooling effectiveness is calculated downstream. The dimensionless temperature contour overlaid with streamlines concluded that the SLD trench hole arrangement with forward injection forms a developed effusion layer due to counter-rotating vortex pairs, which helps in proper mixing of coolant jets into the mainstream and improves film cooling effectiveness in lateral as well as in longitudinal direction. It is observed that super-long-diamond arrangement with forward injection provides the highest film cooling effectiveness than square-diamond and long-diamond arrangements and favours early development of the coolant film layer.

The complex vortex structure compressor leads to the problem that the transition model is insufficient in predicting the flow instability of the compressor. In this paper, the rectangular cascade of compressor of different turning-angle conditions is taken as the object, and the transition characteristics on the end wall and the blade surface of the compressor cascade are in comparison by the method of large eddy simulation/LES. The effects of the horseshoe vortex and the separation bubble over the compressor cascade on the transition process are emphatically discussed. By analyzing characteristic parameters of the vortex structure, it is found that the separated transitional flow corresponds to multiple separations-and reattachments of the shedding vortex, and is affected by the cross-flow transition and the separate-transition. Finally, by discussing the instability of the separation line, reattachment line and the cross-flow inflection point of the separated transitional flow, it reveals that the transient disturbance caused by the vortex motion is an important reason affecting the prediction accuracy of the transition model.