Band 41, Heft 3

Understanding the occurrence of various feedback mechanisms of an under-expanded impinging supersonic jet is a crucial task in research. The presence of several jet modes is examined in this study for the flat and corrugated impinging plate geometries. The behavior of impinging plate configurations during mode switching is investigated by varying the flow state, such as the jet Mach number. The staging behavior at various jet Mach numbers is observed using acoustic spectral plots and schlieren flow visualization. To explore the presence of various types of modes during the jet impingement due to the modification of jet Mach number, ensemble averaging and Proper Orthogonal Decomposition of schlieren images are carried out. In the majority of situations, the corrugated design shows a reduction in tonal noise and overall sound pressure level. In exceptional cases, for the corrugated plates, the enhanced overall sound pressure level is caused by the existence of axisymmetric instability (A1, A2).

The flow performance of a dual wall-mounted cavity in a strut-injector scramjet combustor in steady reacting flow conditions is computationally analyzed. A baseline configuration corresponding to DLR experiments and two proposed configurations with varying bottom wall cavity depth and fixed top wall ramp is considered. Steady-flow computations are performed using the 2-D Reynolds Averaged Navier–Stokes method with k-ω SST turbulence closure coupled and single-step reaction chemistry. The calculated flow patterns, density, pressure, and temperature fields are compared with shadowgraph and wall pressure measurements from DLR experiments. The cavity and strut are mounted downstream of the strut to analyze the shock patterns and their interference with the shear layer mixing features. The estimated flow patterns, density, pressure, and temperature fields are compared with shadowgraph and wall pressure measurements from DLR experiments. Incorporating cavity and ramp configuration provides earlier complete combustion compared to the baseline model, with a marginal rise in the total pressure caused by additional shock wave formation that emanates from the corners of the cavity and ramp. The combustion zone widens in the lateral direction as the cavity shifts the shock train downstream of the strut injector owing to intense shock shear layer interactions.

There are developed methods for high-pressure turbine (HPT) blade loads and remaining useful life (RUL) prediction; however, they are ineffective and time-consuming for in-service HPT blades under actual operating conditions. Hence, it is necessary to use an acceptable computational effort to predict the HPT blade’s load and in-service lifetime. Drawing from the idea of the usage-based life (UBL) prediction method, this paper first proposes a framework for the life digital twin (LDT) to characterize and track the in-service life consumption of the HPT blades under actual operating conditions. The second work mainly focuses on developing the steady-state and transient load calculation surrogate models for the HPT blade’s LDT. Using the developed surrogate models, it can now calculate the steady-state and transient loads of the HPT blade in an acceptable time with the necessary accuracy. The proposed approach is demonstrated on an HPT blade of a typical commercial turbofan engine. Because the operating load of the HPT blade severely affects its in-service lifetime, the application of this approach enables the construction of an LDT of the HPT blade. It can reduce the uncertainty and variability associated with the in-service life prediction of the HPT blade under actual operating conditions.

This paper presents the numerical analysis of a convergent-divergent circular nozzle with the exit Mach number of 1.69 with and without passive control at the exit. The passive control method opted for this analysis was inward and outward ascending triangular protrusion. This paper explores the influence of the passive control geometry and its blockage area concerning the nozzle exit. The nozzle pressure ratio (NPR) used for carrying out the flow analysis were 3, 4.932, and 6. Two different inward and outward protrusions were used with a height of 1.5 mm and 3 mm. From the results, the potential core length of the protrusion 1.5 mm height was not much changed in the both outward and inward cases. But when the height of the protrusion was increased to 3 mm, there was a noticeable core length reduction at all NPR but with different cases. At the NPR of 6, the potential core length of the inward protrusions 3 mm was reduced by 44 % compared to the plain CD nozzle.

As turbine operating conditions change, the blade height and tip clearance undergo continuous alterations due to the combined effects of thermal stress, aerodynamic forces and centrifugal forces, subsequently influencing the turbine performance. To take this effect into account in turbine performance prediction, this study considers the influence of fluid-heat-structure coupling on blade height and tip clearance and establishes a one-dimensional comprehensive prediction method for multi-stage axial turbine performance considering blade height. When compared with experimental results from a four-stage axial turbine, by considering the fluid-thermal-solid coupling effects, the average relative error in total pressure ratio prediction is reduced from 3.76 % to 1.99 % and the average relative error in total temperature ratio prediction is reduced from 2.03 % to 1.26 %. Compared with the traditional flow prediction method, the prediction results of turbine characteristics considering blade height and tip clearance changes in this paper are closer to the experimental results.

In order to explore the air ingestion characteristics of the open-ends squeeze film dampers (SFD), a 3D computational fluid dynamics solution model of the SFD was proposed based. The setting of opening boundary conditions was considered, which was verified experimentally by the bidirectional excitation. The results revealed that with the same precession frequency, the smaller radial clearance leads to the greater oil film damping. An increasing oil supply and oil holes led to the smaller air ingestion range at the end. The design guideline for the open-ends SFD was investigated in the present work.

The article presents the results of an analytical study of the flow dynamics in a microdiffuser. An expression for the velocity profile is obtained with account for slip effects. The effect of the Knudsen number on the velocity profiles is shown. The effect of the microdiffuser angle on the flow velocity profile and the friction coefficient is analyzed. The opening angle of the microdiffuser was determined, at which slip effects do not affect the velocity profile.

In order to realize the most economical operation of helicopter power system, the optimal rotor speed control method is proposed and designed. Firstly, the simplified performance calculation model of helicopter required power and turboshaft engine are established, which constitute the performance calculation model of helicopter power system together. Then, according to the above model, through applying blade loading constraint and selecting the minimum engine fuel flow as the optimization objective, an integrated optimization method of optimal rotor speed based on golden section method is proposed, which reveals the variation law of optimal rotor speed under different operation conditions. Finally, the numerical simulation and hardware-in-loop (HIL) simulation are conducted separately to validate the optimal rotor speed control method. The results indicate that compared with constant rotor speed operation, optimal rotor speed control method can decrease engine fuel flow by more than 21 %, which proves the significant effectiveness and satisfactory feasibility. The optimal rotor speed control method ensures that the helicopter power system operates at the optimal rotor speed, and is beneficial to reach the most economical cruise target.

This manuscript presents Computational Fluid Dynamics simulation of a ramjet combustor using Reynolds Averaged Navier–Stokes approach for turbulence with flamelet combustion modelling. In the approach a one-dimensional premixed flame simulation is performed to cover a wide range of mixture fraction space. After a mesh study, effects of turbulence, Realizable k - ε and SST k - ω , and combustion modelling, Steady Laminar Flamelet and Flamelet Generated Manifold, are investigated in detail. The Flamelet Generated Manifold model that offers a reaction delay by employing an additional progress variable, shows there are no reactions taking place in the vicinity of jet inlets, hence the unrealistic temperature rise that is seen in the case of using Steady Laminar Flamelet is not observed. A study on the choice of progress variable based on the combustion products is carried out. The present study readily demonstrates good predictability of Flamelet Generated Manifold combustion modelling in such a dual inlet ramjet combustion chamber.

A design scheme of a static shaft turbofan engine is proposed to meet the requirements of light weight and large thrust weight ratio of small aeroengine. As the core component of stable combustion, the thermal protection problem of the mid-mounted combustion chamber is particularly prominent. This paper designs a mid-mounted combustion chamber configuration that combines gas film cooling and central combustion. The influence of structural parameters on combustion characteristics is explored by numerical simulation, and the theoretical design and numerical simulation is verified based on combustion test results. The results show that the flame shape of the mid-mounted combustion chamber conforms to the characteristics of central combustion. The combustion effect of the nozzle with spray angle of 45° and flow rate of 1.87 kg/h meets the requirements of secondary combustion of the static shaft turbofan engine, and the air inlet of the combustion liner effectively increases the thickness of the cooling gas film. The experimental results are in good agreement with the numerical simulation results, and the temperature of the combustion liner wall can be reduced effectively. The above research provides a theoretical basis for the combustion chamber design of small static shaft turbofan engines and a reference for the thermal protection methods of small aeroengine combustion chambers.

Both the aerodynamic and thermal performance are important for a high-pressure turbine design. This paper investigates aerothermal effects of opening on the suction side squealer of a cavity tip in a turbine cascade. There are four cases investigated, ‘Cavity’, ‘Opening 1’, ‘Opening 2’ and ‘Opening 3’. For Opening 1 with an opening area at 15 % axial chord, the opening outflow affects the near-tip flow mainly by interacting with the passage vortex, and the near tip loss increases by less than 3 %. For Opening 2 with an opening area at 12 % axial chord, the near tip loss is lower than Opening 1. By further rounding the edge of the opening area, the loss of the tip with opening is less than 1 % higher than the cavity tip. Nevertheless, the opening tip can achieve a reduction of the thermal load of the suction side squealer inner wall by up to 19 %.

The turbine inlet air temperatures exhibit a significant degree of elevation, hence leading to potential adverse consequences such as the degradation of blade material integrity. Consequently, the necessity to cool the turbine blades has arisen, leading to the implementation of various cooling systems. This study aims to conduct a comparative analysis of three different types of pin fins, namely oblong, circular, and teardrop, in a wedge duct. The range of Reynolds number considered for the analysis is between 10,000 and 70,000, whereas the pin fins and endwalls are subjected to a uniform heat flux of 3280 W/m 2 . The findings suggest that the friction factor associated with teardrop pin fins is 28.4 % lesser than circular pin fins and when compared to oblong pin fins; it is reduced by 34.5 %. The findings suggest that the friction factor associated with teardrop pin fins is 14 % lower compared to oblong and circular pin fins. The TPF improves with Re, and it is 24.5 % higher than oblong pin fin geometry and 39.2 % higher than circular shaped pin fins.

A hybrid drone called a biplane quadrotor operates in both low (during the horizontal flight) and high (transition maneuver) Angle of Attack (AoA). So, this paper focuses on enhancing aerodynamic force during the transition maneuver. Synthetic Jet actuators (SJAs) can modify airfoil shapes virtually, so either the flow reattaches or flow separation will be delayed. This delay can enhance the aerodynamic force. In this paper, CFD analysis is performed using Ansys Fluent to study the impact of SJA on NACA 0015 airfoil at high (30°) AoA. This study aims to find the best location of SJA for high AoA to get maximum lift enhancement. The outcome of this study reveals that at 85 %, we can get maximum enhancement in the lift.

This paper conducts a study on closed-loop control of engine performance parameters during mode transition process of TBCC engine based on artificial intelligence method. Firstly, a composite modeling method based on stepwise regression analysis and batch normalization-depth neural network is proposed to establish the on-board model during mode transition to estimate the thrust and inlet airflow in real-time. Secondly, based on the hybrid penalty function-particle swarm optimization algorithm, a mode transition control schedule applicable to the closed-loop control of thrust and inlet airflow is developed. Finally, a data processing method based on similarity conversion is proposed to extend the applicable envelope range of the mode transition control system. The transition time is shortened by 33.3 %, and the fluctuations of thrust and inlet airflow are reduced by 1.33 % and 10.77 %, respectively. When the control system is applied to the off-design mode transition process, a satisfactory mode transition performance is also obtained.

Experiments have been carried out to investigate the over expanded free jet at Mach number 2.0 without and with tabs of different cross-sectional shapes, tab length penetration into the jet core defined as blockage ratio, and symmetrical and asymmetrical placement of tabs at the jet exit. The cross-sectional shapes such as triangular, square, and circular have been studied with the penetration depth of 3 %, 7 %, and 11 % blockage. The Pitot tube and Schlieren flow visualization were carried out in experiments. The basic features of the jet with Mach disk, shock cell, and its distortion with the adoption of tabs at various nozzle pressure ratios indicate a definite influence that helps in the reduction of the core jet length. This characteristic is also influenced by changes in tab cross-section and length of penetration. Since the basic jet has three-dimensional structures, the asymmetric placement of tabs is beneficial compared to the symmetric orientation of tabs. The jet width increases with adoption of these tabs and different shapes. Whereas, a maximum core jet length reduction of the order of 60 % could be achieved using the present techniques.

The controllable speed casing is a novel casing treatment approach that makes partial casing rotate at adjustable and proper speed to achieve stability expansion. Structural parameters of casing treatment are found to influence the effect of stability expansion by many studies. In this paper, the effect of the ending position of the rotatable ring in controllable speed casing on the tip leakage flow and the stability expansion was studied numerically. The results show that when the rotatable ring rotates at 30 % and 50 % rotor design speed, the controllable speed casing achieves the stability expansion of the compressor rotor no matter where the ending position is. The upstream movement of the ending position decreases the axial pressure gradient in the middle and rear of the tip passage. It pushes the shock wave downstream, which reduces blockage region at tip leading edge. The upstream movement of the ending position contributes to an approximately linear increase in the stable operating margin.

In order to explore the influence of the infrared stealth technology on the performance of the integrated helicopter/engine system, an integrated modeling method of helicopter/engine/infrared suppressor is proposed. Firstly, based on the power calculation model of the helicopter, combined with the high-precision turboshaft engine component-level model, an integrated simulation platform is built, which takes into account the nonlinear characteristics. Then, the aerodynamic characteristics of infrared suppressors under different engine operation states are studied by CFD numerical computation method, and the infrared radiation characteristics are obtained through combining the positive and negative ray tracing method and narrow band model method. Ultimately, the utilization of the power turbine outlet stagnation pressure is employed as the pivotal interface linking the turboshaft engine and the infrared suppressor in order to formulate an integrated model encompassing the helicopter, engine, and infrared suppressor subsystems. The simulation results demonstrate that compared with the conventional exhaust system, the application of the infrared suppressor greatly enhances the stealth performance of the helicopter, but also results in the unexpected decrease in engine output power. Moreover, the specific fuel consumption of the turboshaft engine increase and the compressor surge margin decreases in case of the consistent flight condition.

The effect of nozzle lip thickness and velocity ratio on coaxial subsonic jet mixing, at different Mach numbers, has been studied experimentally and numerically. Decay of coaxial subsonic jets emanating from coaxial nozzles of lip thickness 0.7, 1.7 and 2.65 mm with velocity ratio (VR) from 0.2 to 1.0 at primary jet exit Mach numbers of 0.6, 0.8 and 1.0 has been studied. Free jet without co-flow (VR0) was also studied for comparison. Jet centerline Mach number decay, turbulence and velocity variation in the radial direction are analyzed. The results show that mixing the coaxial jet at a low-velocity ratio is better than a high-velocity ratio, at all Mach numbers of the present study. The nozzle lip thickness has a significant influence on the secondary jet. Mixing of the jet in the presence of VR0.2 coaxial jet is found to be the highest. Characteristic decay of Mach 0.8 and 1.0 jet for lip thickness 1.7 and 2.65 mm is faster than lip thickness 0.7 mm. For a given lip thickness, increasing of velocity ratio is found to retard the mixing between primary and secondary jets.

Swirl cooling can not only increase the area of the heat exchange wall covered by the coolant, but also improve the average heat transfer intensity and uniformity of the target surface. SST k - ω turbulence model is utilized in the conjugate heat transfer numerical simulation. Based on C3X blades, leading edge swirl cooling structure of the corresponding areas are modified. The flow and heat transfer characteristics of swirl cooling are analyzed at different cross-sections and positions. It is found that there exists an optimal aspect ratio and hole spacing to minimize the temperature gradient on the swirl cavity wall. The swirling motion in the swirl cavity can significantly increase the heat transfer coefficient of the wall surface.

Different inlet structures have a significant impact on the internal flow characteristics of a solid-magnesium powder water ramjet engine. Based on the magnesium-water reaction model, a computational fluid dynamics (CFD) method is applied to establish a numerical simulation method for the internal flow field of the engine, and the internal flow characteristics of the engine under different inlet structure conditions are studied. The simulation results show that high-temperature gas can effectively promote the ignition of magnesium powder at the top of the combustion chamber, while accelerating the evaporation of the first inlet water and increasing the combustion rate of magnesium powder. The secondary inlet has the most significant effect on the temperature inside the combustion chamber. When the secondary inlet flow rate increases towards the top of the combustion chamber, it increases the amount of heat absorbed by the evaporating water at the top of the chamber, thereby reducing the temperature at the top of the combustion chamber. However, when the flow rate is low, it results in insufficient oxidizer at the top of the combustion chamber, which is unfavorable for the combustion of magnesium powder.

The wide flight range and high torsional vibration frequency of high-speed helicopters impose stricter criteria for the high-bandwidth control of turboshaft engines. Consequently, research is underway to implement a high-bandwidth control method for turboshaft engines using the linear active disturbance rejection control (LADRC) theory. Initially, the LADRC is designed based on the mathematical model of the integrated helicopter/engine system. To address the challenge of maintaining control quality with varying speed reference commands for the power turbine, an improved LADRC method with tracking differentiators (TD) is developed. Numerical simulations comparing the control effectiveness of LADRC with TD to cascade PID and conventional LADRC methods are conducted. The results demonstrate that the improved LADRC gains have a wider tuning range than the LADRC controller, and the power turbine speed tracking effect of LADRC with TD is optimal. It is more conducive to accomplish high-bandwidth control of turboshaft engine with variable rotational speed.

In this study, a CFD computational model has been established to properly simulate complex correctly expanded supersonic flow generated by a three-dimensional C–D nozzle for the purpose of Thrust Vector Control (TVC) simulations. Jet tabs affixed to the perimeter of the nozzle exit are used to achieve thrust vectoring. This study focuses on the estimation of the force on the tabs due to the jet issued from the C–D nozzle using Ansys 19.2 three-dimensional density based solver, implicit formulation with SST K – ω turbulence model. The force is calculated for different tab configurations based on tab sector angles of 60°, 90°& 120° for the designed Mach number and X / D ratio of 0.7. The nozzles with Mach 1.6, Mach 1.8 and Mach 2.0 designed using SolidWorks are compared. It has been observed that with an increase in sector angle, tab force increased by an average of 70 % for sector angle of 60° to 90° and 36 % for sector angle of 90° to 120°.