Fan Pressure Ratio Research Articles

Coupled Aeropropulsive Design Optimization of an Over-wing Nacelle Configuration

The over-wing nacelle (OWN) configuration has the potential to improve on the conventional under-wing nacelle (UWN) configuration by enabling higher bypass ratios and increasing noise shielding. To explore this potential, we perform coupled aeropropulsive design optimization to study the coupled analysis and the design tradeoffs between aerodynamics and propulsion. We find that the variations in wing shape are not significant when the OWN configuration is optimized for different fan pressure ratios. Optimizing the OWN configuration using the proposed approach results in 3% less shaft power than optimizing the wing and propulsor separately. The OWN has marginally superior performance over the UWN for the same lower fan pressure ratio. However, this low fan pressure ratio may be unattainable for the UWN configuration because of ground clearance constraints, whereas the OWN configuration is not subject to such constraints. We study the optimal OWN placement and show that placing it aft of the trailing edge and away from the wing root results in lower required shaft power. In this study, we perform single-point optimizations to understand the fundamentals of the OWN configuration in terms of aerodynamics and propulsion. However, multiple operating conditions, structural integrity, and inclusion of the pylon geometry should be considered to assess the net benefit of the OWN configuration. Nevertheless, these advancements in aeropropulsive optimization are critical to OWN configuration design and more sustainable aircraft.

Reverse Thrust Aerodynamics of Variable Pitch Fans with Inlet Distortion

Abstract Variable pitch fans can improve the operability of low pressure ratio fan systems by re-pitching the rotor blades. If a variable pitch fan can generate sufficient reverse thrust on landing it also eliminates the need for heavy, cascade-type thrust reversers. However, any reverse thrust generation is impacted by high levels of inlet distortion generated as air is drawn into the exhaust nozzle. This paper uses RANS computations and low-speed rig experiments to explore how a representative inlet distortion from the engine installation affects the aerodynamics and performance of a variable pitch fan operating in reverse thrust mode. The simulations and the experiments show that the distortion from the engine installation leads to a highly three-dimensional flow field with a large recirculation region within the bypass duct. The distortion substantially redistributes the mass flow, thrust and power in the engine. Streamline tracking combined with a power balance analysis reveals highly radial flow within the fan rotor, with almost all the fan power used to drive the recirculating flow in the bypass duct, generating high loss and high total temperature. The recirculation reduces the net mass flow through the engine to around 5% of a uniform inflow case and greatly reduces the effective reverse thrust. With uniform inflow, the net reverse thrust was found to be 35% of the nominal takeoff thrust. With inlet distortion, this was reduced to 20%. This demonstrates the importance of designing and operating variable pitch fan systems to minimise reverse thrust inlet distortion.

Performance and Exergy analysis of TurboJet and TurboFan configurations with Rotating Detonation Combustor

Conventional gas turbine is reaching its saturation level and further improvement in the cycle efficiency is possible through redesigning the cycle. One of the effective methods is to introduce a Pressure Gain Combustor (PGC) in place of the conventional deflagration combustor as PGC contributes to a considerable gain in thermal efficiency. Pulsed Detonation Combustor (PDC) and Rotating Detonation Combustor (RDC) are the two most prevalent combustor designs for pressure gain combustion. The major challenges of integrating PGCs to existing conventional cycles are the highly unsteady exhaust flow from PGC, high exhaust temperatures from PGC, potential back flow from combustor to compressor. As turbines are designed for much steadier flows, this unsteady exhaust flow and high temperature exhaust gas from PGC is challenging for the turbine of the cycle and its design. One of the solutions is to have an ejector between the PGC and turbine in order to reduce the flow fluctuations, flow velocity and flow temperature. In this work, different cycle layouts of Turbojet and Turbofan engines working with RDC are described. Low order RDC and ejector models are used in this paper. As the ejector is being used, the compressed air from the compressor is split for RDC, ejector, turbine blade cooling. The different cases for the compressed air split are also examined. The performance parameters of the engines are evaluated for different compressor pressure ratios, fan pressure ratios, bypass ratios. Finally, an exergy analysis is performed on all components.

Conceptual Design of Layered Distributed Propulsion System to Improve Power-Saving Benefit of Boundary-Layer Ingestion

DPS (distributed propulsion system) utilizing BLI (boundary-layer ingestion) has shown great potential for reducing the power consumption of sustainable AAM (advanced air mobility), such as BWB (blended-wing body) aircraft. However, the ingesting boundary layer makes it easier for flow separation to occur within the S-shaped duct, and the consequent distortion due to flow separation can dramatically reduce the aerodynamic performance of the fan, which offsets the power-saving benefit of BLI. By analyzing the source of power saving and power loss of BLI, this paper presents the LDPS (layered distributed propulsion system) concept, in which the freestream flow and boundary-layer flow are ingested separately to improve the power-saving benefit of BLI. In order to preliminarily verify the feasibility of LDPS, an existing DPS is modified. The design parameters and the system performances of LDPS are studied using a 1D engine model. The results show that there is an optimal ratio of the FPR (fan pressure ratio) for the FSE (freestream engine) to the BLE (boundary-layer engine) that maximizes the PSC (power-saving coefficient) of LDPS. This optimal ratio of FPR for the two fans can be obtained when the exit velocities of FSE and BLE are the same. Under the optimal ratio of FPR for the two fans, the PSC of LDPS is improved by 5.83% compared to conventional DPS.

Impact of Installation on the Performance of an Aero-Engine Exhaust at Wind-Milling Flow Conditions

Abstract This paper presents a numerical investigation of the effect of wing integration on the aerodynamic behavior of a typical large civil aero-engine exhaust system at wind-milling flow conditions. The work is based on the dual stream jet propulsion (DSJP) test rig, as will be tested within the transonic wind tunnel (TWT) located at the aircraft research association (ARA) in the UK. The DSJP rig was designed to measure the impact of the installed pressure field due to the effect of the wing on the aerodynamic performance of separate-jet exhausts. The rig is equipped with the dual separate flow reference nozzle (DSFRN), installed under a swept wing. Computational fluid dynamic simulations were carried out for representative ranges of fan and core nozzle pressure ratios (CNPR) for “engine-out” wind-milling scenarios at end of runway (EOR) takeoff, diversion, and cruise conditions. Analyses were done for both isolated and installed configurations to quantify the impact of the installed pressure field on the fan and core nozzle discharge coefficients. The impact of fan and core nozzle pressure ratios, as well as freestream Mach number and high-lift surfaces on the installed suppression effect, was also evaluated. It is shown that the installed pressure field can reduce the fan nozzle discharge coefficient by up to 16%, relative to the isolated configuration for EOR wind-milling conditions. The results were used to inform the design and setup of the experimental activity which is planned for 2023.

Investigation of overall performance of turbofan engine with pulse detonation combustor in bypass duct

The performance model was established to study the performance of mixed exhaust turbofan engine with pulse detonation combustor (PDC) in the bypass duct. The effects of the total pressure loss of isolator and bypass circulation parameters on the characteristic of PDC and performance of engine were analyzed. Then, the sensitivity analysis of performance parameters was carried out on the influence of the components parameters. The performance of mixed exhaust turbofan engine with PDC in the bypass duct was compared with that of the conventional afterburner turbofan engine at the same design circulation parameters. The results show that to increase the total pressure loss of isolator is beneficial to the improvement of performance because of the improvement of pressure ratio; when fan pressure ratio is constant, with the increasing of bypass ratio, the specific fuel consumption and specific thrust increase; with the increasing of fan pressure ratio, when bypass ratio is 0.2 to 0.4, specific thrust increases first and then decreases. When bypass ratio is 0.4 to 0.5, specific thrust increases first and then remains constant. When bypass ratio is 0.5 to 0.9, specific thrust increases, but the rate of increase reduces. At different bypass ratio, specific fuel consumption decreases with increasing of fan pressure ratio; the performance of engine is more sensitive to the component parameters that directly affect the flow state parameters of the bypass duct; because of supercharging performance of PDC, PDC in the bypass duct only uses part of air flow to organize combustion that can make specific thrust of turbofan engine with PDC in the bypass duct and conventional afterburner turbofan engine are equivalent. Meanwhile, specific fuel consumption can be reduced by 27.7% at design point and 12.8%-26.8% at off-design points.

Www.ijeit.misuratau.edu.ly ISSN 2410- Paper ID: EN A Comparison Study For the Performance of the Mixed and Unmixed Flow Two Spool Turbofan Engines at Variable Parameters

The following paper is to investigate the influence of some operational and design parameters on the performance of a turbofan engine with two spools and for mixed and unmixed flow. Because there are many influential variables on the performance, here a computer program with VISUAL BASIC has been developed to carry out this study. The input variables taken in consideration are the flight speed, compressor pressure ratio, fan pressure ratio, the altitude and the combustion temperature. The output obtained at these variables, which are the thrust specific fuel consumption, specific thrust, propulsive efficiency and thermal efficiency are depicted and discussed for the two types of flow.The results show that there is a great influence of these variables on the performance parameters.

Development of a Predictive Tool for the Parametric Analysis of a Turbofan Engine

Parametric cycle analysis, an on-design engine study, specifies the required design characteristics that optimize engine performance. This study aimed to conduct a parametric analysis of a low-bypass turbofan engine with an afterburner, F100-PW229, and develop a technique for estimating its performance based on data using machine learning and deep learning. Commercially available gas turbine simulation software, GasTurb 14, was used to create a dataset of engine performance response variables and input design parameters. The effects of the Mach number, fan pressure ratio, altitude, turbine entry temperature, and bypass ratio on the specific thrust, propulsive efficiency, specific fuel consumption, and total fuel flow were investigated. Regression learning models and deep neural networks were then programmed on this dataset to predict responses for new input data. In MATLAB, a total of 24 regression models were trained with cross-validation, and the model with the least root mean square error was selected as the final model. The machine learning regression models produced reliable output parameter predictions, with the least root mean square error of 9.076 × 10−5. Among the numerous regression models tested, Gaussian process regression, the quadratic support vector machine, and the wide neural network emerged to be the most successful in predicting turbofan engine performance metrics. A multilayer perceptron model was coded in Python with two hidden layers that accurately predicted the performance parameters. The mean square error value on test data was found to be as low as 0.0046. In comparison to intensive computational simulations, machine learning and deep learning models offer an efficient method for conducting parametric analysis of turbofan engines.

Multi-fidelity and multi-objective aerodynamic short nacelle shape optimisation under different flight conditions

AbstractThroughout the course of a flight mission, a range of aerodynamic conditions, including design-point conditions and off-design conditions, are encountered. As the bypass ratio increases and the fan-pressure ratio decreases to reduce the engine’s specific fuel consumption, the engine diameters increase, which results in an increase in the nacelle weight and overall drag. To reduce its weight and drag, a shorter nacelle with a length-to-diameter ratio $L/D = 0.35$ is investigated. In this study, an adaptive cokriging-based multi-objective optimisation method is applied to the design of a short aero-engine nacelle. Two nacelle performance metrics were employed as the objective functions for the optimisation routine. The cruise drag coefficient is evaluated under cruise conditions, whereas the intake pressure recovery is evaluated under takeoff conditions. The cokriging metamodel are refined using an effective infilling strategy, where high-fidelity samples are infilled via the modified Pareto fitness, and low-fidelity samples are infilled via the Pareto front. By combining parameterised geometry generation, automated mesh generation, numerical simulations, surrogate model construction, Pareto front exploration based on the non-dominated sorting genetic algorithm-II and sample infilling, an integrated multi-objective optimisation framework for short aero-engine nacelles is developed. Two-objective and three-objective test functions are used to validate the effectiveness of the proposed framework. After the optimisation process, a set of non-dominated nacelle designs is obtained with better aerodynamic performance than the original design, demonstrating the effectiveness of the optimisation framework. Compared with the kriging-based optimisation framework, the cokriging-based optimisation framework outperforms the single-fidelity method with a higher hypervolume value at the same number of iteration loops.

Robust aerodynamic optimization and design exploration of a wide-chord transonic fan under geometric and operational uncertainties

Axial compressors are inevitably affected by various uncertain factors in the process of manufacture and operation. These uncertainties obviously lead to reduced efficiency and large performance dispersion. However, researches on uncertainty quantification and robust design of compressors still faces severe difficulties due to the complexity of compressor structure and internal flow. This paper aims to present an automated and effective framework for uncertainty quantification and aerodynamic robustness optimization of axial compressor. The manufacturing error distribution is derived from the measurement data of machined fan blades, and the sparse grid-based polynomial chaos expansion method is used to propagate the uncertain factors and predict the probability density distribution of the fan performance. A novel surrogate model that combines a self-organizing mapping and a back-propagation neural network is constructed to explore and visualize the correlation between uncertainty parameters and performance responses. Robust aerodynamic design optimization is achieved based on the genetic algorithm. The results indicate that the coupled neural network model exhibits good accuracy for uncertain approximate modeling. Compared with the prototype fan, the optimized fan's mean isentropic efficiency and pressure ratio increase by 0.97% and 0.72%, respectively. The standard deviation of isentropic efficiency, pressure ratio, and mass flow rate decrease by 46.3%, 21.4%, and 15.2%, respectively. The present study provides a reference and exploration for uncertainty quantification and robust optimization of advanced refined multi-stage turbomachinery.

Study on performance and mechanisms of a novel integrated model with Power & Thermal Management system and turbofan engine

• A novel integrated model is developed by combining PTMS and engine system. • An off-design heat exchanger submodule without specifying detailed geometry is embedded. • The coupling influences of PTMS and engine are revealed over the entire flight envelope. • The weak cooling performance of PTMS arises in high Mach number and altitude regions. With the improvement of modern aircraft performance, thermal management has become increasingly challenging. The power and thermal management system (PTMS) proposed by Honeywell is an integrated approach to solve this problem. Although the PTMS has inevitable mass and energy interactions with the engine, the existing PTMSs are usually designed based on limited flight states without considering the full engine condition. Therefore, we first propose a novel integrated model combining PTMS and engine based on a semi-precise heat exchanger submodule. Parametric analysis is conducted to find out the impact of the engine and PTMS design variables on the propelling and cooling performance of the integrated system, and energy and exergy analyses are adopted to evaluate the degree of coupled effects between PTMS and engine over the whole flight envelope. Results show that the influence of FAN and HPC pressure ratio on refrigerating performance is more significant than the bypass ratio and inlet temperature of HPT at the design point. The cooling efficiency of PTMS is improved by 4.1% and 1.2%, respectively, as the pressure of HPC and FAN rises by 50%. Besides, the refrigerating performance of PTMS gradually weakens with the increasing flight Mach number, while the cooling efficiency reduces by 50% as the Mach number increases from 0 to 1.8. This trend indicates that the working condition and flight environment of the engine need to be considered in the design process of the PTMS system. As for the control variables of PTMS, the refrigerating efficiency of PTMS decreases by 60% when the pressure ratio of PTMS compressor (PTMS_COMP) changes from -33.3% to 33.3%. Inconsistently, the refrigerating efficiency shows a positive correlation with the heat transfer capacity of the kerosene heat exchanger (PTMS_KEHEX) and primary heat exchanger (PTMS_PRHEX). Meanwhile, the impact of the second heat exchanger (PTMS_SEHEX) on the cooling performance of PTMS can be neglected. Above all, in future PTMS designs, the influences of the pressure ratio of PTMS_COMP, the heat transfer capacity of PTMS_KEHEX and PTMS_PRHEX, as well as the engine control variables need to be considered simultaneously.

A Comparative Performance Analysis of the Novel TurboAux Engine with a Turbojet Engine, and a Low-Bypass Ratio Turbofan Engine with an Afterburner

Presented herein is a comparative performance analysis of a novel turbofan engine with an auxiliary combustion chamber, nicknamed the TurboAux engine, against a turbojet engine, and a low bypass ratio turbofan engine with an afterburner is presented. The TurboAux engine is an adaption of the low-bypass ratio turbofan engine, but with secondary combustion in an auxiliary bypass annular combustion chamber for thrust augmentation. The TurboAux engine is envisioned with the desire to facilitate clean secondary burning of fuel at temperatures higher than in the main combustion chamber with air exiting the low-pressure compressor. The comparative study starts by analyzing the turbojet engine and its performance with and without an afterburner segment attached. In parallel, the conventional turbofan and its mixing counterpart are analyzed, also with and without an afterburner segment. A simple optimization analysis of a conventional turbofan is performed to identify optimal ‘fan’ pressure ratios for a series of low-bypass ratios (0.1 to 1.5). The optimal fan pressure ratios and their corresponding bypass ratios are adapted to demonstrate the comparative performance of the varying configurations of the TurboAux engine. The formulation and results are an attempt to make a case for charter aircrafts and efficient close-air-support aircrafts. The results yielded increased performance in thrust augmentation, but at the cost of a spike in fuel consumption. This trade-off requires more in-depth investigation to further ascertain the TurboAux’s utility.

Predicting the energy and exergy performance of F135 PW100 turbofan engine via deep learning approach

In the present study, the effects of flight-Mach number, flight altitude, fuel types, and intake air temperature on thrust specific fuel consumption, thrust, intake air mass flow rate, thermal and propulsive efficiecies, as well as the exergetic efficiency and the exergy destruction rate in F135 PW100 engine are investigated. Based on the results obtained in the first phase, to model the thermodynamic performance of the aforementioned engine cycle, Flight-Mach number and flight altitude are considered to be 2.5 and 30,000 m, respectively; due to the operational advantage of supersonic flying at high altitude flight conditions, and the higher thrust of hydrogen fuel. Accordingly, in the second phase, taking into account the mentioned flight conditions, an intelligent model has been obtained to predict output parameters (i.e., thrust, thrust specific fuel consumption, and overall exergetic efficiency) using the deep learning method. In the attained deep neural model, the pressure ratio of the high-pressure turbine, fan pressure ratio, turbine inlet temperature, intake air temperature, and bypass ratio are considered input parameters. The provided datasets are randomly divided into two sets: the first one contains 6079 samples for model training and the second set contains 1520 samples for testing. In particular, the Adam optimization algorithm, the cost function of the mean square error, and the active function of rectified linear unit are used to train the network. The results show that the error percentage of the deep neural model is equal to 5.02%, 1.43%, and 2.92% to predict thrust, thrust specific fuel consumption, and overall exergetic efficiency, respectively, which indicates the success of the attained model in estimating the output parameters of the present problem.

Sensitivities of Aircraft Acoustic Metrics to Engine Design Variables for Multidisciplinary Optimization

Aircraft environmental performance metrics including fuel burn, takeoff and landing noise, and gaseous emissions are increasingly driving the design and optimization of modern aircraft engines. Gradient-based methods can significantly reduce the computational resources required during their multidisciplinary optimizations. However, current state-of-the art noise modeling tools do not allow for effective computation of sensitivities of acoustic metrics with respect to engine design and control variables, and therefore cannot be easily integrated into multidisciplinary optimization tools. This paper presents the utility of the python Noise Assessment tool (known as pyNA): an aircraft noise estimation model providing sensitivities of acoustic metrics, enabling multidisciplinary optimization and optimal control of engines for low-noise aircraft. The model is used to assess the trades of environmental performance metrics in the clean-sheet engine design for the NASA supersonic technology concept airplane. For the engine examined, if the fan diameter is allowed to change while technology levels are held constant, the takeoff noise levels and emissions are found to be most sensitive to the design fan pressure ratio. For an engine with a fixed fan diameter, takeoff noise levels are found to be insensitive to design variables; whereas nitrogen oxide emissions and cruise thrust-specific fuel consumption are most sensitive to the engine fan and high-pressure compressor pressure ratios.

Boundary-Layer Ingestion Benefit for the STARC-ABL Concept

Boundary-layer ingestion (BLI) offers the potential for significant fuel burn reduction by exploiting tightly coupled aeropropulsive effects. NASA’s Single-aisle Turboelectric Aircraft with Aft Boundary-Layer propulsion (STARC-ABL) concept employs BLI on an electrically powered tail cone thruster to take advantage of the technology on what is otherwise a conventional airframe. Despite the traditional airframe of this concept, aeropropulsive integration is critical to the BLI propulsor’s performance. Therefore, it is vital to employ tightly coupled aeropropulsive models to design BLI systems. In this work, 3-D RANS simulations are used to model the aerodynamics, and 1-D thermodynamic cycle analyses are used to model the propulsion system. The two models are tightly coupled using NASA’s OpenMDAO framework, enabling efficient design optimization through gradient-based optimization with analytic derivatives. Using this coupled aeropropulsive framework, 18 computational fluid dynamics (CFD)-based aeropropulsive design optimizations are performed to study the power requirements of the BLI configuration and a reference podded configuration where the electric fan ingests freestream air. This study provides the first set of CFD-based performance data for the STARC-ABL concept across the design space of BLI fan size and pressure ratio. The results quantify the power savings through BLI compared to a traditional propulsion system.

Performance and operability of an electrically driven propulsor

This manuscript discusses the operation of an electrically driven fan for a hybrid-electric propulsion system for BAe-146 aircraft. The thrust requirement is fed into an integrated cycle and aerodynamic design tool for the sizing of a ducted fan as one of the main propulsors, podded under the wing as a replacement for a turbofan engine. The electric motor design is initiated with the torque and speed requirements and with the dimensional constraints arising from the driven fan geometry. The fan operation and aerodynamic design are derived by changing the fan pressure ratio and hub-to-tip ratio to obtain a 2-D design space. Surface-mounted permanent magnet electric motor designs are mapped on the 2-D fan design space. The design and operational flexibility of the system is investigated through three scenarios. In the first scenario, the aircraft rate of climb is changed to downsize the electric motor. In the second scenario, the electric motor rated frequency is changed to increase the power density and in the third scenario the electric motor current density is changed for the same purpose. The investigated three scenarios provide design and operational guidelines for reducing the weight of the electric motor for a direct drive application.

Low-Noise Design of Medium-Range Aircraft for Energy Efficient Aviation

Promising low-noise aircraft architectures have been identified over the last few years at DLR. A set of DLR aircraft concepts was selected for further assessment in the context of sustainable and energy-efficient aviation and was established at the TU Braunschweig in 2019, the Cluster of Excellence for Sustainable and Energy-Efficient Aviation (SE2A). Specific Top-Level aircraft requirements were defined by the cluster and the selected DLR aircraft designs were improved with focus on aircraft noise, emissions, and contrail generation. The presented paper specifically addresses the reduction of aviation noise with focus on noise shielding and modifications to the flight performance. This article presents the state of the art of the simulation process at DLR and demonstrates that the novel aircraft concepts can reduce the noise impact by up to 50% in terms of sound exposure level isocontour area while reducing the fuel burn by 6%, respective to a conventional aircraft for the same mission. The study shows that a tube-wing architecture with a top-mounted, forward-swept wing and low fan pressure ratio propulsors installed above the fuselage at the wing junction can yield significant noise shielding at improved low-speed performance and reduce critical fuel burn and emissions.

High fidelity modeling tools for engine liner design and screening of advanced concepts

With aircraft engines trending toward ultra-high bypass ratios, resulting in lower fan pressure ratios, lower fan RPM, and therefore lower blade pass frequency, the aircraft engine liner design space has been dramatically altered. This result is also due to the associated reduction in both the available acoustic treatment area (axial extent) as well as thickness (liner depth). As a consequence, there is current need for novel acoustic liner technologies that are able to meet multiple physical constraints and simultaneously provide enhanced noise attenuation capabilities. In addition, recent advances in additive manufacturing have enabled the consideration of complex liner backing structures that would traditionally be limited to honeycomb cores. This paper provides an overview of engine liner modeling and a description of the key physical mechanisms, with some emphasis on the use of low to high-fidelity tools such as empirical models and commercially available software such as COMSOL, Actran, and PowerFLOW. It is shown that the higher fidelity tools are a critical enabler for the evaluation and construction of future complex liner structures. A systematic study is conducted to predict the acoustic performance of traditional single degree of freedom liners and comparisons are made to experimental data. The effects of grazing flow and bias flow are briefly addressed. Finally, a more advanced structure, a metamaterial, is modeled and the acoustic performance is discussed.

Evaluation of a Regional Aircraft with Boundary Layer Ingestion and Electric-Fan Propulsor

This paper presents a conceptual study of a regional aircraft with a turboelectric propulsion system and boundary-layer ingestion. The aircraft has an additional electric propulsor installed at the aircraft tail cone, which is driven by generators installed on both underwing engines, aiming to ingest the fuselage boundary layer and improve the aircraft overall performance. The proposed configuration is an aircraft reengining of the reference aircraft platform, the Embraer 175-E1, targeting minimizing airframe modifications. Parametric variations of the thrust split ratio as well as the electric fan pressure ratio were performed in order to create a design space of possible solutions for the proposed concept and also to provide insights of the aircraft key variables trends. From that, a feasible and optimized configuration in terms of efficiency was selected and compared to the reference aircraft. As the main conclusion, it was determined that the studied propulsion system has the potential to provide specific air range benefits in the order of 4 to 7%, which may not be enough to justify a new development.

Performance assessment of a solid oxide fuel cell turbine-less jet hybrid engine integrated with a fan and afterburners

Reducing carbon emissions in the field of aviation is important with the rising prosperity of air transport. Electric aircraft play an essential role in the process. Hybrid engines, composed of a solid oxide fuel cell, a fan, a compressor, and an afterburner are proposed. Fuel cells provide energy to the compressor and the fan. Internal duct air and external duct air are mixed at the inlet of the nozzle. Propulsion power is output by the nozzle. In order to study the effects of operating parameters on the performance of the hybrid engine, thermodynamic performance models are built. The main conclusions are as follows: (1) The specific thrust and specific impulse of the engine are dramatically increased with increasing bypass ratios. Meanwhile, fuel cell power rises slowly, which indicates that the engine with fans has more advantages than the one without fans in the respect of thrust-weight ratios. (2) Fuel cell power first increases then decreases with increasing fan pressure ratio under the conditions of constant overall pressure ratios. (3) Increasing afterburner temperature is the most effective way to improve the thrust of the engine. When the temperature is 2217 K, the specific thrust is up to 3085 N/(kg⋅s−1). Meanwhile, the specific impulse and thermal efficiency are 2693 s and 60.5%, respectively.