The rotating stall in a centrifugal compressor with a vaneless diffuser was investigated both experimentally and numerically with focus on the effect of the internal flow field within the impeller on the diffuser stall. Through numerical analysis, the boundary layer separation at the impeller outlet was found to play an important role in the expansion and rotation processes of the diffuser stall. In particular, the expanded boundary layer separation near the hub side at the outlet of the main blade (M.B.) suction surface passage was considered to be the main cause of the expansion and rotation processes. A longitudinal vortex existed at the throat of the M.B. passage, and the mass flow rate in the M.B. passage was significantly reduced by the blockage effect. In addition, the longitudinal vortex induced the rolling up flow near the hub side at the impeller exit. Thus, the boundary layer separation expanded.

Casing treatment is an effective technique in extending stall margin of axial and centrifugal compressor. However, its impacts on the stall behaviour of mixed-flow compressor are still not completely understood until now. To conquer this issue, unsteady full-annulus simulations were conducted to investigate the stall mechanism of a mixed-flow compressor with and without axial slot casing treatment (ASCT). The circumferential propagating speed of spike inception resolved by the numerical approach is 87.1% of the shaft speed, which is identical to the test data. The numerical results confirmed that the mixed-flow compressor fell into rotating stall via spike-type with and without ASCT. The flow structure of the spike inception was investigated at 50% design rotational speed. Instantaneous static pressure traces extracted upstream of the leading edge had shown a classic spiky wave. Furthermore, it was found that with and without ASCT, the mixed-flow compressor stalled through spike with the characteristic of tip leakage spillage at leading edge and tip leakage backflow from trailing edge, which is different from a fraction of the centrifugal compressor. The resultant phenomenon provides corroborating evidence for that unlike in axial-flow compressor, the addition of ASCT does not change the stall characteristics of the mixed-flow compressor. The flow structure that induced spike inception with ASCT is similar to the case with smooth casing. In the throttling process, tip leakage flow vortex had been involved in the formation of tornado vortices, with one end at the suction side, and the other end at the casing-side. The low-pressure region relevant to the downward spike is caused by leading-edge separation vortex or tornado vortex. The high-pressure region relevant to the upward spike is induced by blockage from the passage vortex. These results not only can provide guidance for the design of casing treatment in mixed-flow compressor, but also can pave the way for the stall waring in the highly-loaded compressors of next-generation aeroengines.

A 2 MW gas turbine engine has been developed for the distributed power market. This engine features a 7:1 pressure ratio radial inflow turbine. In this paper, influences of various geometry features are investigated including turbine tip and backface clearances. In addition to the clearances, the effects of the inducer deep scallop and exducer rounded trailing edge are investigated. Finally, geometric features associated with a split rotor (separate inducer and exducer) are studied. These geometry features are investigated numerically using CFD. Part of the numerical results is also compared to experimental data acquired during engine test to validate the CFD results.
Results indicated that for this specific turbine, the influences of the exducer radial tip clearance, inducer axial tip clearance, and even scalloped blade backface clearance all have negligible influences on performance. In all cases, 1% increase in clearance only attributes to approximately a 0.1% lower efficiency. This finding is very different from former published papers with low pressure ratio turbines, indicating different flow physics apply for a turbine with a relatively high-pressure ratio.

The S-shape inlet has been widely used in advanced military aircraft due to their advantage of reducing radar signature. However, the curvature of the inlet usually causes different kinds of intake distortion at the aerodynamic interface plane (AIP). Among them, the swirl distortion has been seriously concerned because of its great impact on the performance and stability of aero-engines. There is still no universal criterion for assessing the stability of compressors in the condition of strong swirl distortion. As an approach of assessing the swirl intensity and pattern, vortex identification method may be used as an auxiliary method for stability analysis. In this paper, numerical and experimental investigations on different S-ducts were carried out. The axial vorticity component and Q criterion were used to analyze the quantitative correlation between geometry and swirl intensity. It was found that there is a relatively strong correlation between the geometry, the axial vorticity component and the Q criterion. The present investigation may provide a quick reconstruction method to model the effect of S-ducts for compressor stability prediction.

The turbulence characteristics of the shaped hole film cooling are very complex. In this study, Large Eddy Simulation (LES) and Reynolds-averaged Navier-Stokes (RANS) are used to study the film cooling of the shaped hole. The time-averaged results are compared with the experimental data in the literature. Because of the eddy-viscosity model, the RANS method roughly deals with the simulation of boundary layer, which leads to a large deviation. The RANS results are compared with the LES results to identify the weaknesses of the Realizable k-ε model in predicting the turbulence characteristics of the shaped hole film cooling. The eddy viscosity hypothesis and the temperature gradient diffusion hypothesis are evaluated using LES data. Furthermore, the turbulence characteristics of the in-hole flow are analysed with the help of the incremental Proper Orthogonal Decomposition (iPOD). The turbulence presents strong anisotropy and some convection structures are induced from the shear zone. 

To improve the efficiency and fidelity of the numerical analysis for cascade flutter, we propose an efficient scale-resolving simulation method dedicated to time-periodic flows by incorporating the harmonic balance approach into the large-eddy simulation. This method combined convergence calculations of the steady-state problem based on the harmonic balance method for periodic components, and the nonlinear time-marching method for turbulent fluctuations. Using the proposed method, deterministic periodic components and stochastic turbulent fluctuations were calculated simultaneously, and the effect of turbulent fluctuations on deterministic periodic components was directly calculated without using turbulence models. In this paper, we explain the algorithm and feature of this simulation technique and present the results of the computation for channel flow excited in the streamwise direction as an analysis example using the proposed method. In order to validate the proposed method, an analysis of sinusoidally pulsating channel flow at the central friction-velocity Reynolds numbers   was conducted, confirming that the amplitude and phase of the mean velocity oscillation computed by the proposed method were in good agreement with those of the conventional LES. The present calculation achieved an order of magnitude improvement in computational efficiency compared to conventional LES.

The plasma actuation (PA) effect on the snow falling flow was investigated using a plasma electrode with weather resistant design and the natural snow wind facility of the Hokkaido University of Science. NACA0015 test blade with chord length c of 300 mm was used. Wind tunnel tests were carried out under the angle of the attack of the blade was fixed at 15 degrees, and the main flow velocity is U=5 m/s. PIV (Particle image velocimetry) measurements were conducted on various PA conditions using natural dry snowflakes as a tracer. When the actuator was driven under the condition of the fundamental frequency of F=50 kHz, and the pulsed modulated frequency f of fc/U=1 and Duty ratio (Ratio of plasma ON time to pulse duration time) =1%, movement of snowflakes was controlled the most effectively tested. It was clarified that the fundamental frequency of PA also affects the control of snow flow. Under snowfall conditions, the weather resistant designed plasma electrode has suffered no damage and operated successfully.

To study the feasibility of using machine learning technology to solve the forward problem (prediction of aerodynamic parameters) and the inverse problem (prediction of geometric parameters) of turbine blades, this paper built a forward problem model based on backpropagation artificial neural networks (BP-ANNs) and an inverse problem model based on radial basis function artificial neural networks (RBF-ANNs). The S2 (a stream surface obtained by extending a radial curve in turbo blades) calculation program was used to generate the dataset for single-stage turbo blades, and the back propagation algorithm was used to train the model. The parameters of five blade sections in a single-stage turbine were selected as inputs of the forward problem model, including stagger angle, inlet geometric angle, outlet geometric angle, wedge angle of leading edge pressure side, wedge angle of leading edge suction side, wedge angle of trailing edge, rear bending angle, and leading edge diameter. The outputs are efficiency, power, mass flow, relative exit Mach number, absolute exit Mach number, relative exit flow angle, absolute exit flow angle and reaction degree, which are eight aerodynamic parameters. The inputs and outputs of the inverse problem model are the opposite of that of the forward problem model. The models can accurately predict the aerodynamic parameters and geometric parameters, and the mean square errors (MSEs) of the forward problem test set and the inverse problem test set are 0.001 and 0.000 35, respectively. This study shows that machine learning technology based on neural networks can be flexibly applied to the design of forward and inverse problems of turbine blades, and the models built by this method have practical application value in regression prediction problems.

To enhance the understanding of design characters, which have prominent influences during the fan blade out event, a simplified geometrical and dynamic analysis method was derived, and a typical 2-shaft high bypass ratio turbofan engine was selected and modeled. Based on analytical deriving and engineering experience learned from the real engine failure case, three determinative impact actions were recognized from the fan blade out process. The transient trajectories of these impact actions were researched in analytical method, and then thickness of acoustic lining, quantity of fan blades and threshold load of structural fuse were analyzed as key design characters. 36 serialized fan blade out transient dynamic simulations were conducted by using the 2-shaft high bypass ratio turbofan engine model within different combinations of the three key design factors. The results from geometrical and dynamic analysis matched mainly well with the results from simulations. Characteristic phenomenon in simulation can be explained theoretically. Five conclusions can be summarized from these results. (1) If thickness fan acoustic lining was thinner, the deviation between simplified analytical calculation and simulation were not outstanding to predict Blade-Casing the first impact time and angular position. (2) An appropriate thickness of acoustic lining could make a lower impact stress of fan casing at the first impact. (3) Different thickness of acoustic linings leaded to two impact modes for blade 2, which were tip impact and root impact. (4) Different impact conditions between blade 1 and blade 2 caused remarkable speed components distinction of blade 1, and leaded to a wide range of transient trajectory of blade 1 during FBO event. (5) Thicker acoustic lining in this research can usually find the porper threshold loads setting, which can give a satisfactory outbound vibration. Two details were raised for further research, which were impact behavior of composite material fan blade and honeycomb and influences of wider FBO threshold load ranges in design cases with thinner acoustic lining.

The supersonic multi-hole probe is an essential test tool for wind tunnel experiments, which is necessary to develop the basic research of improving the measurement accuracy and expanding the application of the probes.
This paper theoretically derived a gas compression factor δ_s~f(p*, p_s, κ, λ) to expand the scope of application of Bernoulli’s equation, and discussed the reliability issues of using this factor to solve the velocity and Mach number of the supersonic flow. The research results show that the calculation method of aerodynamic parameters of the supersonic flow proposed in this paper has credibility within one ten-thousandth of the calculation error compared with the calculation of aerodynamic theory. Compared with the algorithm in this paper and the other three algorithms, the calculation errors of the velocity and Mach number of the supersonic flow and the static pressure ratio before and after the shock are all within the range of one ten-thousandth based on the experimental data of a transonic turbine linear cascade. However, the error of the post-wave Mach number is relatively large. Finally, a universal supersonic multi-hole probe calibration algorithm proposed in this paper is suitable for automated non-opposing measurement. It has generally credible and fully considers the shock wave factor. It will improve the theoretical system of multi-hole probes, and provide theoretical guidance and technical support for the supersonic wind tunnel experiment.

Since the transition from rotating stall to surge in a transonic compressor at high speed is very quick, quite often there is no time to take measures to prevent the surge. Therefore, it is desired to find any rotating stall precursors, of which the occurrence can offer sufficient time for stall or surge prevention. In this study, a series of unsteady flow analyses were performed on a transonic compressor under operating conditions before rotating stall with unsteady results scrutinized to find rotating stall precursors. Particular attention is paid to the spatial modes and time modes of static pressure near the casing and around the blade leading and trailing edges. The results show that the characteristics of the precursor in both spatial and time domains can be used as rotating stall warnings.

Partial surge is a type of instability inception discovered in our previous studies. It has been confirmed that partial surge is localized in the blade hub region, and the flow oscillation it caused will lead to the stall cells in the rotor tip. While since all information about partial surge is obtained from the compressor stage experiments, what will happen to the stall process after the stators are removed is also an issue that worth investigating. So, in this paper, a series of experiments are carried out on the single rotor embedded in the transonic compressor stage with partial surge inception. First, the experimental results under uniform inlet conditions show that, although partial surge appears at high rotor speed in the stage case, it does not occur at any speed in the single rotor case. Then, it is found by numerical simulation that the absence of partial surge may be due to the insufficient rotor hub loading, so an experiment with increased hub loading is carried out, but still fails to trigger partial surge. Finally, the reason why partial surge doesn’t occur in the single rotor is discussed. From these results, it can be concluded that partial surge cannot occur in the single rotor case, and the large-scale corner separation in the stator hub is considered to play an important role in the formation of partial surge.

Tip leakage flow has become one of the major triggers for rotating stall in tip region of high loading transonic compressor rotors. Comparing with active flow control method, it’s wise to use blade tip modification to enlarge the stable operating range of rotor. Therefore, three pressure-side winglets with the maximum width of 2.0, 2.5 and 3.0 times of the baseline rotor, are designed and surrounded the blade tip of NASA rotor 37, and the three new rotors are named as RPW1, RPW2, and RPW3 respectively. The numerical results show that the width of pressure-side winglet has significant influence on the stall margin and the minimum throttling massflow of rotor, while it produces less effect on the choking massflow and the peak efficiency of new rotors. As the width of the pressure-side winglet increases from new rotor RPW1 to RPW3, the strength of leakage massflow has been attenuated dramatically and a reduction of 20% in leakage massflow rate has appeared in the new rotor RPW3. By contrast, the extended blade tip caused by winglet has not introduced much more aerodynamic losses in tip region of rotor, and the new rotors with different width of pressure-side winglet have the similar peak efficiency to the baseline. The new shape of the leakage channel over blade tip which replaces of the static pressure difference near blade tip has dominated the behavior of the leakage flow in tip gap. As both the new aerodynamic boundary and throat in tip gap have reshaped by the low-velocity flow near the solid wall of extended blade tip, the discharging velocity and massflow rate of leakage flow have been suppressed obviously in new rotors. In addition, the increasing inlet axial velocity at the entrance of new rotor has increased slightly as well, which is attributed to the less blockage in the tip region of new rotor. In consideration of the increased inlet axial velocity and the weakened leakage flow, the new rotor presents an appropriately linear increase of the stall margin when the width of pressure-side winglet increases, and has a nearly 15% increase in new rotor RPW3.

Coanda jet flap is an effective flow control technique, which offers pressurized high streamwise velocity to eliminate the boundary layer flow separation and increase the aerodynamic loading of compressor blades. Traditionally, there is only single-jet flap on the blade suction side. A novel Coanda double-jet flap configuration combining the front-jet slot near the blade leading edge and the rear-jet slot near the blade trailing edge is proposed and investigated in this paper. The reference highly loaded compressor profile is the Zierke & Deutsch double-circular-arc airfoil with the diffusion factor of 0.66. Firstly, three types of Coanda jet flap configurations including front-jet, rear-jet and the novel double-jet flaps are designed based on the 2D flow fields in the highly loaded compressor blade passage. The Back Propagation Neural Network (BPNN) combined with the genetic algorithm (GA) is adopted to obtain the optimal geometry for each type of Coanda jet flap configuration. Numerical simulations are then performed to understand the effects of the three optimal Coanda jet flaps on the compressor airfoil performance. Results indicate all the three types of Coanda jet flaps effectively improve the aerodynamic performance of the highly loaded airfoil, and the Coanda double-jet flap behaves best in controlling the boundary layer flow separation. At the inlet flow condition with incidence angle of 5°, the total pressure loss coefficient is reduced by 52.5% and the static pressure rise coefficient is increased by 25.7% with Coanda double-jet flap when the normalized jet mass flow ratio of the front jet and the rear jet is equal to 1.5% and 0.5%, respectively. The impacts of geometric parameters and jet mass flow ratios on the airfoil aerodynamic performance are further analyzed. It is observed that the geometric design parameters of Coanda double-jet flap determine airfoil thickness and jet slot position, which plays the key role in supressing flow separation on the airfoil suction side. Furthermore, there exists an optimal combination of front-jet and rear-jet mass flow ratios to achieve the minimum flow loss at each incidence angle of incoming flow. These results indicate that Coanda double-jet flap combining the adjust of jet mass flow rate varying with the incidence angle of incoming flow would be a promising adaptive flow control technique.

This study examines how the complex flow structure within a gas turbine rotor affects aerodynamic loss. An unshrouded linear turbine cascade was built, and velocity and pressure fields were measured using a 5-hole probe. In order to elucidate the effect of tip clearance, the overall aerodynamic loss was evaluated by varying the tip clearance and examining the total pressure field for each case. The tip clearance was varied from 0% to 4.2% of blade span and the chord length based Reynolds number was fixed at 2×105. For the case without tip clearance, a wake downstream of the blade trailing edge is observed, along with hub and tip passage vortices. These flow structures result in profile loss at the center of the blade span, and passage vortex related losses towards the hub and tip. As the tip clearance increases, a tip leakage vortex is formed, and it becomes stronger and eventually alters the tip passage vortex. Because of the interference of the secondary tip leakage flow with the main flow, the streamwise velocity decreases while the total pressure loss increases significantly by tenfold in the last 30% blade span region towards the tip for the 4.2% tip clearance case. It was additionally observed that the overall aerodynamic loss increases linearly with tip clearance.

The rotor blade height with low hub-tip ratio is relatively longer, and the aerodynamic parameters change drastically from hub to tip. Especially the organization of flow field at hub becomes more difficult. This paper takes a transonic 1.5-stage axial compressor with low hub-tip ratio as the research object. The influence of four types of rotor hub contouring on the performance of transonic rotor and stage is explored through numerical simulation. The three-dimensional numerical simulation results show that different hub contourings have obvious influence on the flow field of transonic compressor rotor and stage, thus affecting the compressor performance. The detailed comparison is conducted at the rotor peak efficiency point for each hub contouring. Compared with the linear hub contouring, the concave hub contouring can improve the flow capacity, improve the rotor working capacity, and increase the flow rate. The flow field near blade root and efficiency of transonic rotor is improved. The convex hub contouring will reduce the mass flow rate, pressure ratio and efficiency of the transonic rotor. Full consideration should be given to the influence of stator flow field by hub contouring.

The combustion of ammonia (NH3) has attracted wide interest in fuel vehicle engines, marine engines, and power generators to mitigate carbon dioxide emissions. Unfortunately, the relatively low laminar flame speed presents a technical barrier for this renewable fuel to be used in practice. This work is concerned with numerical examining the effects of elevating inlet temperature on the laminar burning velocity of NH3/air flames with various contents of dimethyl ether (DME) using 1D freely propagating flame calculations, and to shed light on the flame enhancement mechanism. For this, the mechanism is first validated by comparing the numerical predictions with experimental data. Results show that increasing the inlet temperature has a positive effect on the laminar burning velocity of pure NH3/DME/air flames. It is revealed that elevating inlet temperature contributes to a higher adiabatic flame temperature, which is beneficial to the overall chemical reaction rate. Furthermore, the thermal diffusivity of the binary mixture is observed to increase substantially as well. Further kinetic and sensitivities analyses reveal that the inlet temperature has a minimal effect on the reaction pathway, leading to the relative importance of the dominant chain branching over terminating reaction steps to be varied negligibly. The present work confirms that the flame speed enhancement with increasing inlet temperature is primarily the synergetic result of the thermal and diffusion effects, rather than the chemical effect.

Axial Fuel Staging (AFS) technology is an advanced low-emission combustion method in modern gas turbine, which divides the combustor into two axially arranged combustion zones. For revealing the characteristics of axial staged combustion, an industrial-grade combustor was designed and built. The distribution of temperature and velocity field in the combustor was presented with numerical simulation. And an Atmospheric Combustor Test Rig for axial staged combustion was built. The flow resistance characteristics of the combustor were measured at first. Then the effects of the equivalent ratio and the preheating temperature on the pollutant emission and combustion instability were investigated. The results show that the total pressure recovery coefficient in cold state is always above 98%; starting the secondary combustion at low load can reduce NO emissions by 50%, and can suppress the combustion oscillation amplitude of the combustor. At the design point with Φ=0.62 and preheating temperature=400°C, NO emission and CO emission are 15.68 and 4.22 mg/m³ (@15%O₂).

Combustion with lean premixed and low swirl is an effective way of flame organization. It can improve the flame stability and reduce NOx emission. In this kind of combustion, one of the most important issues is fuel/air premixed characteristics. How the structure parameters influence that issue is figured out through numerical simulation. The structure parameters concerned in the study are as follows. They are shape of blades, number of blades, location and shape of gas jet. The influences of them are analysed with comprehensive consideration of many aspects. With the same light shading rate and stagger angle, the axial swirler with curved blades has worse premixed uniformity and lower pressure loss than the one with straight blades. With the same structure of each blade, the decrease of the quantity of blades does influence the pressure loss, while the quantity of gas jets changes correspondingly. But it has little effect on premixed uniformity in a certain range. However, more blades make contribution to better premixed performance. When the total flow area is the same, the axial and circumferential positions of the fuel jets also greatly influence the premixing process. When the fuel jets are upstream the blades and locate at middle of the vanes, the premixing performance is the best. Meanwhile, the jet direction of the fuel jets is a very important influencing factor of the premixing process. When the fuel jet direction is oblique downward at an angle of 30° to the horizontal, the premixing effect is better than the horizontal outflow, which is better than the oblique upward structure.

This paper investigated the effects of variable jetting nozzle angles on the cross-flow suppression and heat transfer enhancement of swirl cooling in gas turbine leading edge. The swirl chamber with vertical jet nozzles was set as the baseline, and its flow fields and heat transfer characteristics were analyzed by 3D steady state Reynolds-averaged numerical methods to reveal the mechanism of cross-flow weakening the downstream jets and heat transfer. On this basis, the flow structure on different cross sections and heat transfer characteristics of swirl chamber with variable jetting nozzle angels were compared with the baseline swirl chamber. The results indicated that for the baseline swirl chamber the circumferential velocity gradually decreased and the axial velocity gradually increased, and the cross-flow gradually formed. The cross-flow deflected the downstream jets and drawn them to the center of the chamber, thus weakening the heat transfer. For swirl chamber with variable jetting nozzle angles, the air axial velocity is axial upstream, opposite to the mainstream, so that the impact effects of cross-flow on the jets were reduced, and the heat transfer was enhanced. Furthermore, with the increase of axial velocity along the swirl chamber, the jetting nozzle angle also gradually increased, as well as the effect of cross-flow suppression, which formed a relative balance. For all swirl chambers with variable jet nozzle angles, the thermal performance factors were all larger than 1, which indicated the heat transfer was enhanced with less friction increment.

A coupling configuration of double-wall cooling and exterior surface thermal barrier coating (TBC) is one of the most promising thermal protection methods of hot components of modern gas turbine. The combined influences of coating thickness, impingement layout, and cooling air flowrate on the overall thermal performances of such configuration were discussed deeply, to provide the valuable guidance of design. Overall effectiveness measurements were implemented under engine-matched Biot numbers and mainstream-to-coolant temperature ratio. Conjugate heat transfer simulations provided the additional information difficult to be acquired by the measurements. The results indicated that the contribution of TBC is much larger than that of increasing the cooling air amount. The thicker TBC can produce the stronger insulation, while the higher risk of thermal damage of itself. The larger coolant flowrate enlarges the benefit of TBC, while the trend is suppressed by the thick TBC. The constant coating thickness cannot bring to the uniform metal temperature, which can be solved through properly adjusting the backside impingement. The trends in overall effectiveness with TBC’s thickness are independent on the change of internal impingement.

Based on the variable characteristics of the actual operating conditions of the turbine shroud and the purpose of improving the cooling effect of the turbine shroud, this paper builds a test system of the impingement-film cooling shroud with two gas inlet angles (90°, 167°). The effects of film cooling hole arrangement, gas inlet angle, blowing ratio (0.7, 1.0, 1.5, 2.0, 2.5, 3.0) and temperature ratio (1.2, 1.3, 1.4, 1.5, 1.6) on the cooling characteristics of the impingement-film cooling shroud were experimentally studied by infrared temperature measurement technology, especially the effects of gas inlet angle and temperature ratio. The results showed that the film covering effect of the film cooling hole vertical or the same direction of the high-temperature gas incoming flow is better than the film covering effect of the reverse direction with the incoming flow, and the optimal arrangement of film cooling holes can improve the cooling effectiveness of the shroud. Compared with 90° intake gas, the film coverage area on the shroud surface of the 167° intake gas is expanded, and the surface average overall cooling effectiveness is increased by 1.03% to 12.6%. The overall cooling effectiveness of turbine shroud increases with the increase of blowing ratio, which increases the flow rate and pressure of cooling gas, and the corresponding increase rate is between 1.04% and 9.96%. However, the increase in the temperature ratio increases the mainstream heating capacity, resulting in a decrease in the cooling effectiveness of the shroud, with a maximum reduction rate of 11.04%.

This paper presents a detailed and comprehensive multiphysics design process of an 80 kW, 60 000 r/min high-speed permanent magnet machine (HSPMM) for a micro gas turbine application. First, the preliminary design of the HSPMM is carried out according to the mechanical and electromagnetic theory. Afterwards, the influence of carbon fiber sleeve (CFS) thickness, rotor diameter and core length on rotor stress and rotor dynamics is carefully analyzed to obtain the optimal range of rotor diameter and core length. On this basis, the electromagnetic and power loss characteristics are analyzed in detail to obtain the final design scheme. Fluid-solid coupling model is used to calculate the temperature field of the HSPMM to verify the rationality of the scheme. The rotor thermal stress analysis considering the multi-layer and multi-angle winding of CFS is carried out to obtain the rotor models suitable for prototype and mass production, respectively. Finally, the prototypes are manufactured and tested to verify the reliability of the multiphysics design process.

Gas turbines are increasingly and widely used, whose research and production reflect a country’s industrial capacity and level. Due to the changeable working environment, gas turbines usually work under the condition of simultaneous changes of ambient temperature, load and fuel. However, the current researches mainly focus on the change in single condition, and do not fully consider the simultaneous change in different conditions. On the basis of single condition, this paper further studies the dual off-design performance of gas turbines under three conditions: temperature-load, fuel-load and fuel-temperature. Firstly, the whole machine model of a gas turbine is established, in which the compressor model has the greatest impact on the performance of gas turbines. Therefore, this paper obtains a more accurate compressor model by combining the engineering modeling advantages of gPROMs and the powerful mathematical calculation ability of MATLAB neural network. Then, according to the established gas turbine model, the dual off-design performance is studied, which is mainly based on the parameter of output and efficiency. The result shows that the efficiency and power output of gas turbines will decrease with the increase of ambient temperature. With the decrease of fuel calorific value, power output and efficiency will increase. As the load decreases, the efficiency of the gas turbines will decrease, and these changes are consistent with the single off-design performance. However, when the fuel and temperature change simultaneously, only adjusting the IGV angle cannot avoid the surge when the temperature is above 30°C. At this time, it is necessary to adjust the extraction rate in order to ensure the safe and stable operation of gas turbines. Therefore, the research on dual off-design performance of gas turbines has an important significance for the peak shaving operation of gas turbines.