High Bypass Ratio Turbofan Engine Research Articles

Prediction of Crosswind Separation Velocity for Fan and Nacelle Systems Using Body Force Models: Part 1: Fan Body Force Model Generation without Detailed Stage Geometry

Modern aircraft engines must accommodate inflow distortions entering the engines as a consequence of modifying the size, shape, and placement of the engines and/or nacelle to increase propulsive efficiency and reduce aircraft weight and drag. It is important to be able to predict the interactions between the external flow and the fan early in the design process. This is challenging due to computational cost and limited access to detailed fan/engine geometry. In this, the first part of a two part paper, we present a design process that produces a fan gas path and body force model with performance representative of modern high bypass ratio turbofan engines. The target users are those with limited experience in turbomachinery design or limited access to fan geometry. We employ quasi-1D analysis and a series of simplifying assumptions to produce a gas path and the body force model inputs. Using a body force model of the fan enables steady computational fluid dynamics simulations to capture fan–distortion interaction. The approach is verified for the NASA Stage 67 transonic fan. An example of the design process is also included; the model generated is shown to meet the desired fan stagnation pressure ratio and thrust to within 1%.

Numerical analysis of hub cavity leakage

Purpose This paper aims to predict the effect of the hub cavity leakage on the overall performance with numerical simulations in the fan/booster of a high bypass ratio turbofan engine. Design/methodology/approach Simulations are conducted for leakage at the fan, the outlet of guide vane and the three-stage booster, as well as hub leakage (contain cavities and sealing). The results obtained are compared to the corresponding simulations without hub leakage. Findings The rotor/stator interaction locations are evaluated to discover a better location. The results show that the seal tooth structure produces secondary flow and turbulence in the root of blade suction surface, which increases the aerodynamic loss. The sealing clearance should be controlled to shrink the turbulent region and decrease the leakage. Practical implications This work can provide a theoretical guidance and technical support for the compressor design, which avoid many repeated manufactures and reduce waste of resources. Originality/value This work improves the understanding of the impact mechanism of hub cavity leakage on the performance when the clearance size of seal is variable.

The Nonlinear Single Controller of DGEN380 Aero Engine Design

The advanced nonlinear sliding mode control method of DGEN380 aero engine is presented in this paper. This aero engine is a small high bypass ratio turbofan engine by which the nonlinear control approach of the aero engine is invested. And this paper focuses on the power management function of the aero engine control system which includes steady control and transient control. The mathematical model of DGEN380 aero engine is built by a set of nonlinear dynamic equation that is validated by experimental data. The single controller based on sliding mode approach is designed that can keep some certain thrust levels during steady state and maintain repeatable performance during transient operation from one requested thrust level to another. The single controller can offset the impact of the signal noise and harmonic disturbance at a certain power point. And the dynamic performance of the single controller is satisfactory at the transient process. The experiment is conducted by aero engine test bench for the single control.

Investigation of Vibration Characteristics of Titanium Wide-Chord Fan Blade

The vibration characteristics of a wide-chord fan blade affect the safety, reliability, and operability of high-bypass ratio turbofan engines. In this study, the vibration characteristics of a fan blade were investigated using numerical and experimental methods. First, the natural frequencies and mode shapes were obtained and verified. Then, the rotating speed and frequency margins were employed to analyze the stability of the blade. The critical crossings in the Campbell diagram were indicated and evaluated. Next, the strain energy density index and weaklink were used to evaluate the dangerous locations for the high-cycle fatigue characteristics of the blade. Finally, single- and multi-axis random-excitation tests were performed. Results indicated that low-order bending modes were dominant under circumferential- and axial-direction excitations, while the torsional mode response was higher under radial-direction excitation. The weaklinks for different excitation directions exhibited different characteristics. This work presents a comprehensive understanding of the fan-blade vibration and response characteristics.

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

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

Axisymmetric hub-endwall profile optimization for a transonic fan to improve aerodynamic performance based on an integrated design optimization method

This paper investigates an optimization for the axisymmetric hub-endwall profile of a transonic fan designed for a civil high bypass ratio turbofan engine. A time-saving integrated design optimization method is proposed based on a new hub-endwall profile modification method in combination with a novel two-stage polynomial response surface. The profile modification method is proposed to change the profile of the baseline hub curve by moving three control points, so as to redesign the hub-endwall. The two-stage polynomial response surface is proposed to obtain an accurate high-order polynomial while requiring fewer sampling points, so as to reduce the computational cost. The results show that the integrated design optimization method is suitable and valuable. After optimization, the corner separation of the studied fan is weakened, and the stall margin of the core fan is also enlarged. Moreover, the aerodynamic performance of the core fan at the design speed is improved.

Effect of mixing Mach number and mixing efficiency on the preliminary cycle design of mixed high-BPR turbofans

This article presents the implementation of an updated analytical flow mixing model in a state-of-the-art, non-dimensional gas turbine cycle performance simulation and optimisation tool. The model considers three separate streams in the mixer, each expanding through its own ‘virtual’ nozzle. The use of three streams, compared to one single stream, allows for a more realistic simulation of a mixed exhaust gas turbine. This approach is used in a parametric study to assess the effect of the choice of mixing efficiency and mixer inlet Mach number on the preliminary design of mixed-exhaust, high-bypass ratio turbofan engines. It was found that in terms of best thermal performance, a trade-off exists between mixer inlet Mach number and mixer effectiveness. The findings of this research establish some useful guidelines for the accurate selection of these two parameters to achieve robust cycle designs.

First and Second Law Analysis of Radical Intercooling Concepts

An exergy framework was developed taking into consideration a detailed analysis of the heat exchanger (HEX) (intercooler (IC)) component irreversibilities. Moreover, it was further extended to include an adequate formulation for closed systems, e.g., a secondary cycle (SC), moving with the aircraft. Afterward, the proposed framework was employed to study two radical intercooling concepts. The first proposed concept uses already available wetted surfaces, i.e., nacelle surfaces, to reject the core heat and contributes to an overall drag reduction. The second concept uses the rejected core heat to power a secondary organic Rankine cycle and produces useful power to the aircraft-engine system. Both radical concepts are integrated into a high bypass ratio (BPR) turbofan engine, with technology levels assumed to be available by year 2025. A reference intercooled cycle incorporating a HEX in the bypass (BP) duct is established for comparison. Results indicate that the radical intercooling concepts studied in this paper show similar performance levels to the reference cycle. This is mainly due to higher irreversibility rates created during the heat exchange process. A detailed assessment of the irreversibility contributors, including the considered HEXs and SC, is made. A striking strength of the present analysis is the assessment of the component-level irreversibility rate and its contribution to the overall aero-engine losses.

A Computational Study to Investigate the Effect of Altitude on Deteriorated Engine Performance

This study presents an investigation on the effect of operational altitudes on the performance of the deteriorated engine. A two-spool high bypass ratio turbofan engine is used as the test subject for this study. The engine is modelled in Gas Turbine Simulation Program (GSP) based on an existing engine model from literature. Real flight data were used for the validation. Deterioration rate of 0.1% per day is applied for all turbofan components engine. The simulation is performed by varying the altitude from sea level until 9000m. Results obtained show reduction in air mass flow rate and engine thrust as altitude increases. The reduction in air mass flow rate is due to the lower air density at higher altitude hence reduces amount of engine thrust. At 1000m to 4000m, thrust specific fuel consumption (TSFC) of the engine is improved compared to sea level. However depleted in TSFC is shown when the aircraft flies at altitude higher than 4000m. At this altitude, the effect of air density is dominant. As a result, the engine is required to burn more fuel to provide a higher thrust to sustain the aircraft speed. More fuel is consumed hence depletion in TSFC is obtained.

Development of Parametric Modeling System for the Whole High Bypass Ratio Turbofan Engine based on Secondary Development of UG

In this paper, the method of developing parametric design system based upon UG with VB (Visual Basic) and Microsoft Excel has been studied. The development strategy and implement method of the parametric modeling system for the whole high bypass ratio turbofan engine based on secondary development of UG have been put forward. The parametric modeling system for the whole high bypass ratio turbofan engine has been developed, and the parametric solid model of the whole high bypass ratio turbofan engine can be established rapidly with this system. The design period of the high bypass ratio turbofan engine was shortened, the design efficiency was heightened, and the design quantity was improved obviously. The system structure and some program codes were presented in the paper. The method presented in this paper has reference value to developing the parametric design system of complex products.

Civil turbofan engine exhaust aerodynamics: Impact of bypass nozzle after-body design

It is envisaged that the next generation of civil large turbofan engines will be designed for greater bypass ratios when compared to contemporary architectures. The underlying motivation is to reduce specific thrust and improve propulsive efficiency. Concurrently, the aerodynamic performance of the exhaust system is anticipated to play a key role in the success of future engine architectures. The transonic flow topology downstream of the bypass nozzle can be significantly influenced by the after-body geometry. This behavior is further complicated by the existence of the air-flow vent on the nozzle after-body which can have an impact on the performance of the exhaust system. This paper aims to investigate the aerodynamics associated with the geometry of the bypass nozzle after-body and to establish guidelines for the design of separate-jet exhausts with respect to future large turbofan engines. A parametric geometry definition has been derived based on Class-Shape Transformation (CST) functions for the representation of post-nozzle-exit components such as after-bodies, plugs, and air-flow vents. The developed method has been coupled with an automatic mesh generation and a Reynolds Averaged Navier–Stokes (RANS) flow solution method, thus devising an integrated aerodynamic design tool. A cost-effective optimization strategy has been implemented consisting of methods for Design Space Exploration (DSE), Response Surface Modeling (RSM), and Genetic Algorithms (GAs).The combined approach has been deployed to explore the aerodynamic design space associated with the bypass nozzle after-body geometry for a Very High Bypass Ratio (VHBR) turbofan engine with separate-jet exhausts. A detailed investigation has been carried out to expose the transonic flow mechanisms associated with the effect of after-body curvature combined with the impact of the air-flow vent. A set of optimum curved after-body geometries has been obtained, with each subsequently compared against their respective conical representation. The obtained results suggest that no significant performance improvements can be obtained through curving the nozzle after-body relative to the case of a conical design. However, it is shown that the application of surface curvature has the potential to unlock new parts in the design space that allow analysts to reduce the required after-body length without any loss in aerodynamic performance. The developed approach complements the existing tool-set of enabling technologies for the design and optimization of future large aero-engines, consequently leading to increased thrust and reduced Specific Fuel Consumption (SFC).

Simulations of an aircraft with constant and pulsed blowing flow control at the engine/wing junction

In high-lift and especially landing conditions, the vortices emerging from the engine strake and pylon as well as from the inboard and outboard slat tips of civil aircraft configurations can be the cause of the so-called nacelle wake flow separation. This phenomenon might be responsible for the wing stall and sudden lift loss; the effect is expected to increase with the coming up of enlarged nacelles and larger slat cutouts typical of Ultra High Bypass Ratio turbofan engines. In the framework of the European project AFLoNext, DLR and ONERA performed numerical studies focused on a realistic aircraft in high-lift configuration. Structured and unstructured grids were generated to perform Reynolds Averaged Navier–Stokes computations aimed at analyzing the baseline flow without control and determining relevant location and settings for active flow control systems. Constant blowing devices with different slot sizes, types and injection velocities were evaluated with RANS simulations over complete lift polars. The gains in CLmax were quantified; they range from 1 to 3% for actuator capacities compatible with manufacturer requirements, the nacelle wake separation appears at angles of attack one to two degrees higher than without control and the lift levels in post-stall conditions are significantly improved. Finally, unsteady RANS computations were carried out to investigate the potential of a pulsed blowing device. The gains in CLmax and flow separation containment obtained with the latter are similar to those of constant blowing devices which yet exhibit greater momentum coefficient and mass flow rate values.

Aero-Engines Intake: A Review and Case Study

This paper presents an aerodynamic review for intakes of aero-engines. Classification of different types is introduced for both subsonic and supersonic aircrafts. Important parameters influencing intake performance like pressure recovery and spillage drag are discussed. In addition, crucial performance issues such as the angle of attack, ground vortex, swirl flow, noise propagation and icing are presented. These are critical issues which are carefully considered by airframe rather than engine manufacturers. Intake icing may be extremely dangerous and leads to drastic decrease in air mass flow and consequently aircraft propulsion. Calculation of intake performance is explained. The effect of intake performance on the whole engine performance is highlighted. Numerical analyses for present and previous researches are explained and critical parameters are identified. Finally, a case study resembling the intake of a high bypass ratio turbofan engine (in close similarity to that of GE CF-6 Turbofan Engine) is numerically analysed using the commercial code “ESICFD 2010” is thoroughly described and analysed. This paper is the first of a series of research work that will cover several R & D activities in HBPR and future EBPR turbofan intakes.

Numerical Prediction of the Influence of Thrust Reverser on Aeroengine’s Aerodynamic Stability

AbstractA numerical method was developed to predict the aerodynamic stability of a high bypass ratio turbofan engine, at the landing stage of a large transport aircraft, when the thrust reverser was deployed. 3D CFD simulation and 2D aeroengine aerodynamic stability analysis code were performed in this work, the former is to achieve distortion coefficient for the analysis of engine stability. The 3D CFD simulation was divided into two steps, the single engine calculation and the integrated aircraft and engine calculation. Results of the CFD simulation show that with the decreasing of relative wind Mach number, the engine inlet will suffer more severe flow distortion. The total pressure and total temperature distortion coefficients at the inlet of the engines were obtained from the results of the numerical simulation. Then an aeroengine aerodynamic stability analysis program was used to quantitatively analyze the aerodynamic stability of the high bypass ratio turbofan engine. The results of the stability analysis show that the engine can work stably, when the reverser flow is re-ingested. But the anti-distortion ability of the booster is weaker than that of the fan and high pressure compressor. It is a weak link of engine stability.

Forced and Aeroelastic Responses of Bird-Damaged Fan Blades: A Comparison and Its Implications

The aeroelastic response of a bird-damaged fan stage at the inlet of a high-bypass ratio turbofan engine is examined using a combined computational fluid dynamics and computational structural dynamics framework. The damaged fan sector consists of five blades obtained from accurate numerical simulation of the bird impact. Forced and aeroelastic response calculations are performed and compared to assess the role of aeroelastic coupling. The calculations are performed at 100, 75, and 60% throttle settings to investigate the role of engine speed on the fan response. Results from the forced response and aeroelastic response calculations indicate that the undamaged blades opposite the damaged sector exhibit the highest level of structural response. Comparing the forced response with the aeroelastic response shows increased participation of the higher structural modes, especially for the damaged blades, that grow in time or exhibit beating. Examination of the work performed by the aerodynamic forces suggests that the growth in blade response is due to aeroelastic phenomena and can cause a potential instability. The results illustrate the importance of aeroelastic effects when predicting the post-bird-strike fan response.

New adaptive PID control method based on IFA-ELM for aero-engine

For the strong nonlinear and variable parameters properties of high bypass ratio turbofan engine,an adaptive PID control method based on optimized extreme learning machine( ELM) was proposed to predict the engine's parameters. To improve the prediction accuracy and the real-time property of ELM,an improved Firefly algorithm( IFA) for multi-peak optimization was adopted to optimize the network parameters of the ELM,and formed an optimized ELM training method IFA-ELM. Under the premise of ensuring prediction accuracy,the algorithm effectively simplified the network scale and improved its generalization capability.The engine fan speed prediction model was built by this algorithm,and gradient descent method was adopted to adjust the PID parameters online based on the model to improve the dynamic performance of engine. Digital simulation results show that compared with conventional PID control,IFA-ELM based adaptive PID method shortens the settling time by 0. 2 ~ 1. 4 s,and reduces the overshoot by 0. 2% ~ 1. 5%,which demonstrates the effectiveness of the proposed control method.

Prediction of the Broadband Noise Radiated from a Fan Rotor of a High-Bypass Ratio Turbofan Engine: Take-off Condition

The purpose of this study is to predict the broadband noise generated by the fan rotor of a high bypass ratio turbofan engine at take-off condition. A numerical investigation of the unsteady aerodynamics of a fan rotor was described, with emphasis on acoustics. Three case studies were examined; namely, design point, a choked point and a near surge point. It was found that the dominant broadband noise mechanisms are due to interaction of turbulence of the incoming flow with the engine casing and nose, interaction of the turbulent boundary layers on the rotor blades with their trailing edges, and interaction of the rotating blade with the turbulence in the incoming flow. The most important noise source was the interaction of the turbulent boundary layers on the rotor blades with their trailing edges. Choked case scored the highest noise level as it has the maximum turbulent kinetic energy levels, maximum mass flow rate and thus maximum relative Mach number. Near-surge case has the lowest noise level except at the lower part of the suction side of the rotor where the reverse flow and vortices caused higher levels of noise.

Erosion of an Axial Transonic Fan due to Dust Ingestion

This paper deals with the prediction of the particle dynamic and erosion characteristics due to dust ingestion in an axial flow fan, installed in a high bypass-ratio turbofan engine that operates in a dusty environment. Dynamic behavior comprises the particle trajectory and its impact velocity and location. While the erosion characteristics are resembled by the impact frequency, erosion rate, erosion parameter and the penetration rate. The study was carried out in two flight regimes, namely, takeoff, where the sand particles are prevailing, and cruise, where the fly ashes are dominated. In both cases, the effect of the particle size on its trajectory, impact location, and the erosion characteristics was studied. To simulate the problem in a more realistic manner, a Rosin Rambler particle diameter distribution was assumed at takeoff and cruise conditions. At takeoff, this distribution varies from 50 to 300 μm with a mean diameter of 150 μm sand particles. While at cruise, this distribution varies from 5 to 30 μm with a mean diameter of 15 μm fly ash particles. The computational domain employed was a periodic sector through both the fan and its intake bounding an angle of (360/38) where the number of fan blades is (38). The intake is a stationary domain while the fan is a rotating one and the FLUENT solver is used to solve this problem. Firstly, the flow field was solved in the computational domain using the Navier-Stokes finite- volume supported by the Spalart-Allmaras turbulence model. The governing equations, representing the particle motion through the moving stream of a compressible flow are introduced herein to calculate the particle trajectory. The solution of these equations is carried out based on the Lagrangian approach. Next, empirical equations representing the particle impact characteristics with the walls are introduced to calculate the rebound velocity, the erosion rate, erosion parameter, impact frequency and penetration rate. Moreover, a method to smoothen the irregularity in the calculated scattered data was discussed as well. During takeoff flight regime, the pressure side of fan blade experienced higher particle impact and erosion damage. The highest erosion rate was found at the corner formed by blade tip and trailing edge of pressure side. During cruise conditions, less erosion rates resulted. Maximum erosion rates are found at the leading edge of the pressure side.

Study of Jet Noise Reduction on Separated Exhaust System Using Chevron Nozzles

The jet noise reduction of chevron nozzles was investigated on high bypass ratio turbofan engine separated exhaust system using both computational predictions and scale model experiments. Six different exhaust nozzles are designed including one baseline nozzle and five different chevron nozzles. The jet noise experiments were carried out in the anechoic chamber. Tam and Auriault’s jet noise prediction theory and MGBK theory were used to predict the noise spectra of different exhaust nozzles. The results show that the far-field noise spectra as well as the noise reduction benefits of chevrons are predicted correctly by the two theories although some discrepancies occur at the high frequency range, and Tam and Auriault’s jet noise theory can give relatively more accurate prediction results. chevron nozzles reduce jet noise at the low frequencies, but increase it at high frequencies.

Unsteady aerodynamics and aeroacoustics of a fan rotor of a high-bypass ratio turbofan engine

The purpose of this study is to predict the broadband noise generated by the fan rotor of a high bypass ratio turbofan engine. A numerical investigation of the unsteady aerodynamics of a fan rotor is described, with emphasis on acoustics. Two flight conditions are discussed: cruise and takeoff. For each case, three case studies are examined: namely, operating point, a choked point and a near surge point. It is found that the dominant broadband noise mechanisms are due to interaction of turbulence of the incoming flow with the engine casing and nose, interaction of the turbulent boundary layers on the rotor blades with their trailing edges and interaction of the rotating blade with the turbulence in the incoming flow. The most important noise source is the interaction of the turbulent boundary layers on the rotor blades with their trailing edges. Choked case exhibits the highest noise level as it has the maximum turbulent kinetic energy levels, maximum mass flow rate and thus maximum relative Mach number. Near-surge case has the lowest noise level except at the lower part of the suction side of the rotor where the reverse flow and vortices causes higher levels of noise.