High Bypass Ratio Turbofan Research Articles

Novel Nuclear Power and Fossil Fuel Hybrid Propulsion Systems for Long-Endurance Unmanned Aerial Vehicles: Configuration Comparison and Performance Optimization

Aircraft nuclear propulsion system can significantly enhance the endurance, but its large size and weight become disadvantages against the conventional fossil fuel propulsion system. In this paper, a novel nuclear power and fossil fuel hybrid propulsion system, which combines the lithium-cooled fast reactor and the kerosene-fueled turbine engine, is proposed to meet the power requirements of different flight segments. Performance assessment and comparison among different hybrid configurations is performed. Results indicate that the configuration combining the nuclear power module with a high bypass ratio turbofan engine is the most promising for aircraft. The primary optimization shows that there is a trade-off between the weight of the nuclear power module and the specific thrust, and the optimal weight of nuclear power module is 3942 kg, with the specific thrust of 195.8 N·kg-1·s-1. The aircraft equipped with the hybrid propulsion system operates at the maximum altitude of approximately 18.2 km and reaches a maximum Mach number of 0.71, allowing for sustained flight for several months in cruise segment. The turbofan engine, nuclear reactor, ancillary systems, shield, and heat exchanger account for 11%, 23%, 11%, 17% and 1% of the fuel mass of the original unmodified aircraft.

On Hysteresis In A Variable Pitch Fan Transitioning To Reverse Thrust Mode And Back

Abstract A novel hysteresis phenomenon during the transition to and back from the reverse thrust mode in a Variable Pitch Fan (VPF) is identified and characterised in this work. This is done by using a three-dimensional (3D) fully transient Unsteady Reynolds averaged Navier?Stokes (URANS) with the transitioning fan blade aerofoils simulated by an adaptation of the mesh displacement method. The VPF is modelled to be transitioning in a modern 40000 lbf geared high bypass ratio turbofan engine architecture at "Approach Idle" engine power setting in a typical twin-engine airframe with the flaps, slats, and spoilers set for an aircraft touchdown airspeed of 140 knots. The transition to reverse thrust mode involves flow starvation into the engine, formation of recirculation zones in the bypass duct and the establishment of the reverse stream, all of which occurs in the opposing presence of the free stream flow at aircraft touchdown velocity. The transition back to forward flow mode involves the gradual re-establishment of the free stream which is opposed by the presence of the reverse stream within the engine. It is quantified that in the transition to reverse thrust, the blockage develops with a larger time delay than the disappearance of the blockage during the transition back due to the interplay of the temporal dynamics of fan blade motion and flow field response. The hysteresis phenomena described in this work are critical in properly developing control schedules to adapt for potential bi-stable flow field development during the landing run.

A Validated Unsteady Turbofan Engine Performance Simulation Using Pseudo Bond Graph Approach

Abstract In this paper, the unsteady performance of a turbofan engine is illustrated and modeled using the pseudo-bond graph approach. The pseudo-bond graph method allows multiphysical and interdisciplinary systems to be described analytically and comprehensibly across disciplines. In this study, a high bypass ratio turbofan engine is modeled using a pseudo-bond graph model that can be adapted for hybrid electric propulsion systems. The development of the model is followed by a full-scale test run of the in-house V2500-A1 turbofan engine. The highly instrumented turbofan engine enables the time-resolved recording of performance-specific variables, thus facilitating the validation of the numerical model. Slam acceleration and deceleration are performed to analyze the unsteady performance. It is shown that the pseudo-bond graph model is able to predict the dynamics of the turbofan engine. High deviations are explained by the incomplete modeling of the control system.

Bayesian Inference for the Calibration of Progressive Damage Model of Dovetail Specimens from Laminated Composite Fan Blade

Abstract Composite fan blades are the preferred alternative for the fan stage of most advanced high bypass ratio turbofan engines. The ability of the dovetail to safely bear this load is one of the key points of the multi-level “test pyramid” approach of compliance demonstration for special airworthiness regulation. Debonding between adjacent layers is the main damage mode of laminated composite fan blades. However, there is difficulty in measuring the as manufactured interlaminar mechanical properties used in finite element models. In this study, tensile loading was applied to simulate the interacting centrifugal force and capture mixed mode damage evolution. Structural responses and material damages were calibrated with measured tensile loads through Bayesian inversion. Posterior probability distributions of maximum interface tractions and contact stresses were solved using Markov chain Monte Carlo sampler. Results indicated the two bilinear cohesive material models had a capacity of predicting empirical means of longitudinal reaction forces as that in test considering additional discrepancy term (0.035 kN and 0.96 kN respectively), while they made an significant impact on the prediction of tensile load history especially when two delamination cracks initiated and propagated. Interface elements provided a higher matching quality in predicting loading history and capturing damage mechanism in association with in-plane progressive damage analysis. This calibrated parameter set could be functioned as benchmark in numerically determining the ultimate tensile load of dovetail elements and reducing the necessary number of physical tests at elemental length level.

Actuator fault diagnosis and severity identification of turbofan engines for steady-state and dynamic conditions

Actuator faults can be critical in turbofan engines as they can lead to stall, surge, loss of thrust and failure of speed control. Thus, fault diagnosis of gas turbine actuators has attracted considerable attention, from both academia and industry. However, the extensive literature that exists on this topic does not address identifying the severity of actuator faults and focuses mainly on actuator fault detection and isolation. In addition, previous studies of actuator fault identification have not dealt with multiple concurrent faults in real time, especially when these are accompanied by sudden failures under dynamic conditions. This study develops component-level models for fault identification in four typical actuators used in high-bypass ratio turbofan engines under both dynamic and steady-state conditions and these are then integrated with the engine performance model developed by the authors. The research results reported here present a novel method of quantifying actuator faults using dynamic effect compensation. The maximum error for each actuator is less than 0.06% and 0.07%, with average computational time of less than 0.0058 s and 0.0086 s for steady-state and transient cases, respectively. These results confirm that the proposed method can accurately and efficiently identify concurrent actuator fault for an engine operating under either transient or steady-state conditions, even in the case of a sudden malfunction. The research results demonstrate the potential benefit to emergency response capabilities by introducing this method of monitoring the health of aero engines.

Design and Numerical Analysis of an Annular Combustion Chamber

Designing a combustion chamber for gas turbines is considered both a science and an art. This study presents a comprehensive methodology for designing an annular combustion chamber tailored to the operating conditions of a CFM-56 engine, a widely used high bypass ratio turbofan engine. The design process involved calculating the basic criteria and dimensions for the casing, liner, diffuser, and swirl, followed by an analysis of the cooling sections of the liner. Numerical simulations using NUMECA software and the HEXPRESS meshing tool were conducted to predict the combustion chamber’s behavior and performance, employing the κ-ε turbulence model and the Flamelet combustion model. Methane was used as the fuel, and simulations were performed for three fuel injection angles: axial, 45°, and 60°. Results demonstrate that the combustion chamber is properly dimensioned and achieves complete combustion for all configurations. The pressure ratio is 0.96, exceeding the minimum design criteria. Additionally, the emissions of unburned hydrocarbons are zero, while nitrogen oxides and carbon monoxide levels are below regulatory limits. These findings validate the proposed design methodology, ensuring efficient and environmentally compliant combustion chamber performance.

Low-frequency sound attenuation by coiled-up meta-liner with nonuniform cross sections under grazing flow

We report a kind of coiled-up meta-liners with nonuniform cross sections (CMNC), which can efficiently attenuate low-frequency sound waves under grazing flow with a deep subwavelength thickness (e.g., ∼λ/17 at 500 Hz). At a grazing flow Mach number of 0.26, the average transmission loss of the meta-liner is 12.6 dB at 500–1000 Hz, which is twice as much as that of a double-degree-of-freedom acoustic liner of the same size. Physically, the nonuniform cross-sectional distribution and significant cross-sectional area ratio enhances vortex shedding, thus resulting in severe acoustic energy dissipation. The excellent low-frequency acoustic attenuation performance of CMNC is investigated thoroughly with experimental, theoretical, and numerical methods. This work provides an avenue for low-frequency noise reduction in grazing flow scenarios (e.g., in a high bypass ratio turbofan engine).

Air-Ground Characteristic Test and Analysis on Fan Blade Vibration Stress of Turbofan Engine

Obtaining the vibration stress of fan rotor blades under actual working conditions through flight tests is a necessary means to carry out design verification and optimization improvement of turbofan engine. Taking the high bypass ratio turbofan engine as research object, installed ground and flight tests were carried out. The vibration characteristics and stress distribution of the fan blade were obtained by direct measurement means of strain gauge, and the air-ground characteristics of the test data were compared and analyzed. The results show that under the steady state, the vibration stress level of fan blade in ground test is significantly higher than that in flight test; There is a significant difference in frequency distribution of excitation factors on fan blade in the ground and flight tests.

Development trends and future expectations of jet engine

Nowadays, nearly all large commercial aircraft are powered by gas turbine engines. Having been developed for about 80 years of engines applied on commercial aircraft, from low bypass ratio turbojet engines to high bypass ratio turbofan engines since the first turbojet engine was patented in 1930 by Frank Whittle, gas turbine engines still have a high potential to improve the performance. Currently, motor thermodynamic efficiency can still rise in the following decades; composite materials, for instance, gamma titanium aluminum and polymer matrix composites, are becoming widely used in airliner propulsion due to their lightweight and high-temperature capability. Accordingly, environmental issues related to gas emissions have also become vital due to the continued flourishing of commercial airliners. In this paper, the development of gas turbine engines applied on commercial airliners from low bypass ratio turbojet engine "Nene1" to high bypass ratio turbofan engine "GE9X," and the current advancement of jet engines applied will be discussed in order to gain possible improvement on the performance of jet engines. Current conditions and expectations of jet engines will be discussed with specific examples, focusing on efficiency, materials and manufacturing, and environmental issues. In the paper, the engines mentioned are all related to commercial aircraft.

A versatile volume-based modeling technique of distributed local quadratic convergence for aeroengines

For advanced aero-engine design and research, modeling and simulation in a digital environment is indispensable, especially for engines of complicated configurations, such as variable cycle engines (VCE) and adaptive cycle engines (ACE). Also, in the research of future smart engines, reliable real-time digital twins are paramount. However, the 2 dominant methods that used in solving the simulation models, Newton-Raphson (N-R) method and volume-based method, are not fully qualified for the study requirements, because neither of them reaches the satisfactory balance of convergence rate and calculating efficiency. In this study, by deeply analyzing the mathematical principle of these 2 methods, a novel modeling and solving method for aero-engine simulation, which integrates the advantages of both N-R and volume-based methods, is established. It has distributed architecture and local quadratic convergence rate. And a novel modeling method for variable area bypass injectors (VABI) is put forward. These facilitate simulation of various configurations of aero-engines. The modeling cases, including a high bypass-ratio (BPR) turbofan and an ACE, illustrate that the novel technique decreases the iterations by about two-thirds comparing with volume-based method, while the success rate of convergence remains over 99%. This proves its superiority in both convergence and calculating efficiency over the conventional ones. This technique can be used in advanced gas turbine engine design and control strategy optimization, and study of digital twins.

Designing of non circular air intakes for subsonic gas-turbine engines

The subject matter of the article is the process of subsonic air intake shaping for gas-turbine engines at the airplane preliminarily design stage. The goal is to develop a mathematical model for non-circular air intake shaping for gas-turbine engines on the base of V. I. Polikovskii method of subsonic air intake shaping for high-bypass ratio turbofan. The tasks to be solved are: to consider the possibility of non-circular shape of the external outline of the engine nacelle; to take into account the possibility of non-circular shape of the internal air intake duct (in the first approximation, the shape of internal air intake duct cross-section is defined in the form of a rectangular with possible four different radiuses in its corners); to consider the engine inlet spinner presence. The methods used are: analytical and digital mathematical methods, implemented in MathCAD and Microsoft Visual Studio systems. The following results were obtained: On the base of the proposed method, new calculation module for the Power Unit software version 11.8 has been developed (С-language Win32 UNICODE application) having a friendly user interface. Conclusions. The scientific novelty of the results obtained is as follows: 1) mathematical model (algorithm and its program implementation) for non-circular air intake shaping for gas-turbine engines has been developed considering non-circular shape of the external outline of the engine nacelle, non-circular shape of the air intake duct internal outline, presence of engine inlet spinner, and zero expansion angle in the diffuser outlet cross-section; 2) adequacy of calculation results by the developed mathematical model is shown by means of comparison with the shape of real air intake, developed by the Antonov Company. For the following improvement of the mathematical model, it is desirable to add the possibility of considering S‑shape of the air intake duct, defining its length from designer’s considerations, defining a bigger radius of curvature of the air intake lip, and considering the presence of boundary layer bleeding devices in front of the air intake.

Model reduction of a cyclic symmetric structure exhibiting geometric nonlinearity with a normal form approach

This papers deals with the computationally costly simulation of the frequency response function of geometrically nonlinear structures exhibiting cyclic symmetry. Its objective is to propose a reduction method that extends the model of a single sector reduced via the direct normal form to a full cyclically symmetric system. After validation on simple structures, the methodology is applied on a 3D meshed blade, representative component of the new generation of high bypass ratio turbofans. Based on both the proposed reduction methodology and theoretical conditions that predict the possible appearance of internal resonances as a function of the number of sectors and the engine order of the excitation, the complex nonlinear dynamics of a finely meshed full fan cyclic system is then investigated. This could not have been achieved without an appropriate reduction method. Branch switching algorithm, stability analysis, and bifurcation tracking compose the numerical tools employed along the harmonic balance method to simulate the forced response of the structure.

Nacelle intake flow separation reduction at cruise condition using active flow control

Turbofan engine intakes are designed to provide separation-free flow at the fan face over a wide range of operating conditions. But at some off-design conditions, like at high flight speeds and high angles of attack (AoA), the aero engine intake may encounter flow separation. This boundary layer separation inside the nacelle inlet of an aircraft engine can lead to a large number of undesirable outcomes like reduction in fan efficiency, engine stall and high levels of stress on the fan blades. Active flow control is a promising solution to reduce inlet boundary layer separation and the associated fan-face flow distortion at such off-design conditions. By blowing pressurized air into the intake near the separation point, the boundary layer is energized and separation can be controlled. This study investigates the applicability of lip blowing, an active flow control technique, to control intake separation and flow distortion at the fan-face. First, intake separation was triggered in a 3D CFD model based on the NASA Common Research Model (CRM) using high AoA cases at cruise condition (Mach number 0.85, Mass flow capture ratio ∼0.7) and the features of separated flow were analyzed. Thereafter, active flow control was introduce to the intake in the form of two types of lip blowing, direct and pitched blowing. The efficacy of lip blowing at achieving separation control in an ultra high bypass ratio turbofan engine intake has been established through this study. The present paper also examines the significance of blowing parameters like the type of blowing, blowing pressure ratio, and blowing slot dimension, at different angles of attack to identify the critical control parameters. Our research successfully establishes proof of concept by demonstrating the feasibility of using lip blowing for separation control in aero-intakes, via numerical modelling. Furthermore, this study also provides crucial insights regarding the important variables to be considered for future experimental studies, and also for detailed studies covering a wider range of operating and blowing conditions.

Design of circular air intakes for subsonic turbofans

The subject matter of this article is the process of subsonic air intake shaping for high-bypass ratio turbofan at the airplane preliminarily designing stage. The goal was to improve a mathematical model of V. I. Polikovskii method of subsonic air intake shaping for high-bypass ratio turbofan. The tasks are to consider the presence of cant of inlet cross-section, required to perform effective operation at airplane cruising angle-of-attacks; to increase the radius of curvature of the air intake lip to provide air flow near it without flow separation, which was definitely determined and could not be increased in the existing method; to improve constant length velocity gradient law (used in this method) so that too large duct expansion angles near the air intake outlet cross-section can be avoided; to consider the engine inlet spinner presence. The methods used are analytical and digital mathematical methods, implemented in MathCAD and Microsoft Visual Studio systems. The following results were obtained: based on the proposed method, new calculation module for the Power Unit software version 11.8 has been developed (С-language Win32 UNICODE application) with a friendly user interface. Conclusions. The scientific novelty of the results obtained is as follows: 1) mathematical model (algorithm and its program implementation) for circular turbofan air intake shaping has been improved considering cant of the inlet cross-section, air intake lip rounding with two radiuses, presence of engine inlet spinner, and zero expansion angles in the diffuser outlet cross-section; 2) adequacy of calculation results using the improved mathematical model is shown using comparison with shapes of circular turbofan air intakes, developed by the leading aviation companies.

On the effects of a separation bubble on fan noise

A scale model of a fan stage operating at approach condition is investigated using wall-resolved Large Eddy Simulations. The configuration is a state-of-the-art fan stage (rotor + stator) of an Ultra High Bypass Ratio turbofan engine, and the computational cost has been reduced by simulating a radial slice of a periodic sector. At low fan speeds and Reynolds number, a recirculation bubble can be observed on the suction side of the fan blades, near the leading edge. This recirculation bubble causes the boundary layer transition to turbulence and is associated with high levels of wall pressure fluctuations. The noise spectra upstream of the fan blades show high frequency tones, which are related to the noise generation mechanism of the recirculation bubble. It has been found that the size, position and pressure disturbances in the bubble depend on the mass flow rate. Furthermore, the frequencies and amplitudes of the high frequency tones in the noise spectra also depend on the mass flow rate. As the mass flow rate decreases, the size of the bubble increases, the bubble shifts towards downstream locations, and high levels of wall pressure fluctuations can be found along the suction side of the fan blade. The acoustic signature of the bubble shifts towards lower frequencies with the mass flow rate, but its amplitude increases. A coherence analysis and a Dynamic Mode Tracking technique are used to confirm that high frequency tones can be generated by the recirculation bubble in the separation region. The flow features at these frequencies suggest that a vortex shedding mechanism is related to the noise radiated from the recirculation bubble.

Impact of three-dimensional turbulence modelling on rotor/stator interaction broadband noise

An accurate prediction of Fan/Outlet-Guide-Vane (OGV) interaction broadband noise (BBN) is fundamental to correctly characterise the noise footprint of high bypass ratio turbofan engines. A three-dimensional synthetic turbulence BBN prediction methodology, which accounts for cascade effects and the airfoil geometry using a frequency-domain Linearised Navier–Stokes Solver, is presented in this work. The methodology computes the BBN generated by an infinite three-dimensional (3D) linear cascade using a spectral representation in the spanwise direction. The 3D problem is reconstructed analytically using two-dimensional (2D) simulations, including additional source terms coming from the spanwise spatial derivatives, leading to a high computational efficiency compared to a standard 3D approach. The methodology has been applied to assess the impact of oblique gusts on a Fan/OGV and a low-pressure turbine airfoil, and the results are compared to a 2D methodology and a previously proposed quasi-3D (q3D) formulation. Differences of more than 10 dB arise when comparing the 3D results with a purely 2D approach caused by the over-estimation of the turbulent kinetic energy that can produce cut-on acoustic modes. The validity of a q3D correction to mitigate this 2D effect has been assessed, demonstrating differences smaller than 2 dB concerning the 3D formulation, and therefore, its use can be justified for rapid predictions. The 3D methodology removes the unphysical impact of the 2D cascade resonances that appear in the q3D approach. The differences between the 3D and q3D approaches have been assessed in detail using an azimuthal noise decomposition. Even though the acoustic response can be remarkably affected by changes in the spanwise wavenumber, the noise azimuthal decomposition and frequency spectra exhibit a similar shape because the acoustic response is relatively constant within a significant fraction of the spanwise wavenumbers that enable sound radiation.

Performance and flow evolution of windmilling utilizing a combination of semi-empirical speed and CFD models during mode transition of the wide-chord fan

With reductions in the pressure ratio of the fan and increases in bypass ratio of the turbofan in recent years, the engine windmilling and relight capability have attracted more and more attention in aviation safety. This paper proposes a combination of the speed model and CFD to examine the internal flow evolution of the wide-chord fan under the windmilling condition. Firstly, we establish a semi-empirical speed model for predicting the fan rotational speed of a high-bypass ratio turbofan at any flight Mach number and altitude based on the flow characteristics of a windmilling fan. The model can provide an initial windmilling airflow at the specified speed for subsequent CFD calculations to reduce the number of iterations. Then, a three-dimensional CFD model is proposed to obtain the operating boundary of the windmilling fan according to the direction of energy transfer between the fan and the airflow. The mechanisms of fan internal flow evolution and tip leakage flow when it operates at a constant speed line were analyzed. The results show that the work distribution profiles along the blade span exhibited self-similarity. The fan blade operated in both spanwise and chordwise mixed configuration. The flow moved gradually from compressor-like regions near the hub to turbine-like regions near the shroud, where the stirrer mode was the transitional stage between them during windmilling. When the rotor entered the windmilling, it operated at a significantly off-design angle of attack, that caused a large area of flow separation to occur near the pressure surface (PS) instead of on the suction surface (SS). The flow separation and loss gradually increased from the hub to the shroud. Meanwhile, the direction of tip leakage flow also changed, a streamwise vortex core was observed on the PS near the blade tip. The core mixed with the low-energy fluid on the PS, resulting in flow blockage, that changed the working profile of the rotor. As the mass flow rate increased to the turbine operating point, the intensity of interaction and influence area between the leakage flow and the main flow further increased, and the aerodynamic loss increased significantly.

Influence Factors for Impact Actions and Transient Trajectories of Fan Blades after Fan Blade Out in Typical 2-Shaft High Bypass Ratio Turbofan Engine

Influence Factors for Impact Actions and Transient Trajectories of Fan Blades after Fan Blade Out in Typical 2-Shaft High Bypass Ratio Turbofan Engine

Empirical estimation of engine-integration noise for high bypass ra-tio turbofan engines

To investigate the impact of installation on jet noise from modern high-bypass-ratio turbofan engines, a model-scale noise experiment with a jet propulsion system and a fuselage model in scale was conducted in the anechoic wind tunnel of ONERA, CEPRA 19. Two area ratios (an area of the secondary nozzle over an area of the primary nozzle), 5 and 7, and various airframe configurations such as wing positions relative to the tip of the engine nacelle and flap angles, were considered. Based on the analysis of experimental data, an empirical model for the prediction of engine installation noise was proposed. The model comprises two components: one is the interaction be-tween the jet and the pressure side of the wing, and the other is the interaction between the jet and the flap tip. The interaction between the jet and the pressure side of the wing contributes to the noise at the low frequencies (≤ 1.5 kHz), and the interaction between the jet and the flap tip con-tributes to the noise at the high frequencies. The proposed model showed a good agreement with the experimental data.

An acoustic liner design methodology based on a statistical source model

The attenuation of fan tones remains an important aspect of fan noise reduction for high bypass ratio turbofan engines. However, as fan design considerations have evolved, the simultaneous reduction of broadband fan noise levels has gained interest. Advanced manufacturing techniques have also opened new possibilities for the practical implementation of broadband liner concepts. To effectively address these elements, practical acoustic liner design methodologies must provide the capability to efficiently predict the acoustic benefits of novel liner configurations. This paper describes such a methodology to design and evaluate multiple candidate liner configurations using realistic, three dimensional geometries for which minimal source information is available. The development of the design methodology has been guided by a series of studies culminating in the design and flight test of a low drag, broadband inlet liner. The excellent component and system noise benefits obtained in this test demonstrate the effectiveness of the broadband liner design process. They also illustrate the value of the approach in concurrently evaluating multiple liner designs and their application to various locations within the aircraft engine nacelle. Thus, the design methodology may be utilized with increased confidence to investigate novel liner configurations in future design studies.