Limit protection, which frequently exists as an auxiliary part in control systems, is not the primary motive of control but is a necessary guarantee of safety. As in the case of aircraft engine control, the main objective is to provide the desired thrust based on the position of the throttle; nevertheless, limit protection is indispensable to keep the engine operating within limits. There are plenty of candidates that can be applied to design the regulators for limit protection. PID control with gain-scheduling technique has been used for decades in the aerospace industry. This classic approach suggests linearizing the original nonlinear model at different power-level points, developing PID controllers correspondingly, and then scheduling the linear time-invariant (LTI) controllers according to system states. Sliding mode control (SMC) is well-known with mature theories and numerous successful applications. With the one-sided convergence property, SMC is especially suitable for limit protection tasks. In the case of aircraft engine control, SMC regulators have been developed to supplant traditional linear regulators, where SMC can strictly keep relevant outputs within their limits and improve the control performance. In aircraft engine control field, we all know that the plant is a nonlinear system. However, the present design of the sliding controller is carried out with linear models, which severely restricts the valid scope of the controller. Even if the gain scheduling technique is adopted, the stability of the whole systems cannot be theoretically proved. Research of linear parameter varying (LPV) system throws light on a class of nonlinear control problems. In present works, we propose a controller design method based on the LPV model to solve the engines control problem and achieve considerable effectiveness. In this paper, we discuss the design of a sliding controller for limit protection task of aircraft engines, the plant of which is described as an LPV system instead of LTI models. We define the sliding surface as tracking errors and, with the aid of vertex property, present the stability analysis of the closed-loop system on the sliding surface. An SMC law is designed to guarantee that the closed-loop system is globally attracted to the sliding surface. Hot day (ISA+30° C) takeoff simulations based on a reliable turbofan model are presented, which test the proposed method for temperature protection and verify its stability and effectiveness.

The conversion of heat into mechanical shaft power or electrical energy is still the worldwide most used electric and propulsion energy source. Whether coal, liquid hydrocarbons, natural gas, nuclear or solar heat are used as primary energy source such cycles are continuously improved aiming at better efficiency, at flexible operating conditions or at reducing manufacturing or maintenance cost. The classic thermal power cycle design uses detailed simulation of the cycle fluid’s changes of state (typically represented by averaged pressure and temperature or specific enthalpy) along the flow paths. Components of such a cycle may be ducts, filters, mixers, compressors, combustors, heat exchangers, nozzles, turbine stages etc. The notations of exergy and anergy were invented in the nineteen-fifties but they still play an astonishingly silent role in education. Knowing already the Carnot factor a student can understand the second law of thermodynamics by the simple fact that the exergy part of energy can be converted into any other form of energy while the anergy part cannot be converted into any exergy form of energy. The most used approach via the saying “that the state variable entropy can only grow” is for most students a mental detour because they do not understand the tie of energy conversion and entropy. The exergy loss breakdown of a cycle gives an instructive view to opportunities and challenges. This paper explains methods to calculate a breakdown of the exergy losses with the example of a typical but notional turbofan using commercially confirmed component performance data. The considerations cover exergy losses by heat exchange, by combustion, by heat loss to ambient, by mixing different fluids, by internal pressure losses, by friction and dissipation in turbomachinery and by the bottoming heat discharge as well as in the propulsive process. Quantified exergy losses for such effects make the limitations of technical improvements visible without going into detailed design work. The purpose of this paper is to motivate for teaching and using exergy-based considerations especially in the basic performance simulation. Additionally the context is occasion for an extra plea on replacing the colloquial expression “energy consumption” by “exergy consumption”. The reader of this paper should have basic knowledge of thermodynamic cycle performance simulation methods and of turbomachinery.

Due to the low level of profile losses already reached in the design of modern low-pressure turbines for turbofan applications, a renewed interest is devoted to the other sources of loss, and namely to the secondary losses. At the same time, the importance of secondary losses has been reinforced by the current design trend towards high-lift profiles. A great attention, therefore, is dedicated to reliable and effective prediction methods as well as on the correct understanding of the mechanisms that drive the secondary flows. In this context, a systematic numerical and experimental campaign was carried out focusing on the impact of different inlet boundary layer (BL) profiles and considering a state-of-the-art low-pressure turbine cascade. Starting from a computational environment representative of a design standard, detailed RANS analyses were carried out in order to establish dependable guidelines for the computational setup. As a major result, such analyses also underlined the importance of the shape of the inlet BL very close to the endwall, hence suggesting tight requirements for the characterization of the experimental environment. The impact of the inlet BL profile on the secondary flow development was experimentally investigated by varying the profile shape very close to the endwall as well as on the external part with respect to a reference condition. The effects on the cascade performance were evaluated focusing on the intensity of the over-under-turning as well as on the associated losses (intensity and penetration) by measuring the span-wise distributions of flow angle and total pressure losses at the cascade exit plane. For all the inlet conditions, comparisons between CFD and experimental results are discussed. Besides providing guidelines for a proper numerical and experimental setup, the present paper underlines the importance of a detailed characterization of the inlet BL for an accurate assessment of the secondary flows. From a broader perspective, when aiming at reproducing (numerically or experimentally) a real engine environment, this suggests that an in-depth matching of the inlet profiles is crucial for reliable estimates of the secondary losses.

Modern trend in installation design is moving towards very high-bypass ratio turbofans. Very high-bypass turbofans represent an effective way of improving the propulsive efficiency of civil aero-engines. Such engines require larger and heavier nacelles, which partially offset the gains in specific fuel consumption. The penalty associated with a larger installation can be mitigated by adopting thinner walls for the nacelle and by shortening the intake section. Such inlet sections are characterized by more restrictive operation condition because they are more prone to separation at high incidence flight conditions. Moreover, in short nacelle installations the by-pass guide vanes and pylon are closer to the fan blades and consequently the distortion due to potential effects induced by the presence of the pylon and non-axisymmetric OGV stage play a significant role in terms of unsteady interaction in the entire system. It is mandatory to consider the inlet, fan, bypass and pylon as a unique coupled system also at the design stage, for assessment of fan force. This kind of assessment is usually carried on by expensive URANS calculation. The factors leading to high computational demands are the spatial resolution required in the fan domain and the time resolution required to sample the fan blade passing frequency. Large savings are therefore possible if simplifications are introduced which relax the resolution requirements in the fan passages and change the nature of the computation into a steady-state computation for the ducts. The present contribution documents a simplified fan model for fan-intake computations based on the solution of the double linearization problem for unsteady, transonic flow past a cascade of thin aerofoils with finite mean load. The coupling with the intake flow and the bypass is performed by using the flow patterns at fan face and fan exit as boundary conditions for the fan model and computing circumferentially non-uniform boundary conditions for the intake and the bypass from the fan model. The computation of the flow in the intake, bypass and pylon is therefore reduced to a steady problem, whereas the computation of the flow in the fan is reduced to one steady problem and a set of linearised models in the frequency domain. The model is applied to a well-documented test case and compares favourably with experimental data and much more expensive three-dimensional, time domain computations.

In the context of an engine design calculation, isentropic or polytropic efficiencies of turbomachinery components are assumed at the outset of the cycle analysis and their values are updated or validated following the aerodynamic design of the components. In the present paper, aerodynamic design calculations of axial-flow compressors and turbines are directly integrated into the corresponding performance component models. This creates a consistent, single-step preliminary design and performance modelling process using a relatively small number of physical and geometric inputs. The aerodynamic design for establishing a component’s overall efficiency is accomplished through a mean-line, stage-by-stage approach where the stagewise isentropic efficiency is calculated employing either loss or semi-empirical correlations. From this process, the stagewise flow annulus radii are also obtained and are used to axially size the component stages assuming the blade aspect ratio and axial gapping distributions. The component flowpath geometry is then produced by simply “stacking” axially the component stages. The developed method is validated against publicly available data for a high-pressure compressor and a low-pressure turbine. Finally, the effectiveness of the method is demonstrated by considering the multi-point design of a High Bypass Ratio Geared Turbofan Engine with bypass Variable Area Nozzle.

The effects of high shaft power offtake in a direct drive, a geared drive, and a novel turbofan configuration are investigated. A design and off-design performance analysis shows the configuration specific limitations and advantages. The more electric aircraft (MEA) concept promises to offer advantages with respect to aircraft performance, maintenance and operating costs. The engines for the MEA concept are based on conventional turbofan architectures. These engines are designed for significantly increased shaft power offtake that is required by the airframe, and the shaft power is usually taken off the high-pressure spool. This can impair the off-design performance of the engine and lead to compromises during engine design and to operability limitations. Taking the power off the low-pressure spool mitigates some of the problems but has other limitations. In this work, an alternative novel turbofan architecture is investigated for its potential to avoid the problems related to high shaft power offtakes. This architecture is called the dual drive booster because it uses a summation gearbox to drive the booster from both the low- and high-pressure spool. The shaft power, if taken off the booster spool, is effectively provided by both the low- and high-pressure spools, which allows the provision of very high power levels. This new concept is benchmarked against a two-spool direct drive and a geared drive turbofan. Furthermore, it is described, how the new architecture can incorporate an embedded motor generator. The presented concept mitigates some of the problems which are encountered during high power offtake in conventional configurations. In particular, the core compressors are less affected by a change in shaft power offtake. This allows higher power offtakes and gives more flexibility during engine design and operation. Additionally, the potential to use the new configuration as a gas turbine-electric hybrid engine is assessed, where electrical power boost is applied during critical flight phases. The ability to convert additional shaft power is compared with conventional configurations. Here, the new configuration also shows superior behavior because the core compressors are significantly less affected by power input than in conventional configurations. The spool speed and its variation is more suitable for electrical machines than in conventional configuration with low-pressure spool power transfer. The dual drive booster concept is particularly suited for applications with high shaft power offtakes and inputs, and should be considered for propulsion of more electric aircrafts.

This paper describes an experimental study on the acoustic performance when mixer nozzles were applied to the core of a subscale turbofan engine. The primary concern of the mixer nozzle is how to satisfy both less jet mixing noise emission and minimum impact on engine performance parameters such as thrust and fuel consumption. A notched nozzle, a nozzle with tiny dents on the trailing edge, initiates small disturbances in the shear layer, weakens the shear stress, and suppresses jet noise. The Japan Aerospace Exploration Agency (JAXA) and IHI Corporation have studied notched nozzles and found that finer and more notches are preferable for both acoustic and aerodynamic performance. As a next step, it is necessary to maintain the tradeoff between noise suppression and impact on engine performance. To evaluate both the acoustic and aerodynamic performance with the notched nozzle, a subscale turbofan engine, DGEN380, was adopted as a demonstration engine. Experiments with this engine were conducted both in a test cell and in an open test site. The notched nozzle, together with a baseline conical nozzle and a referential serrated nozzle, i.e., a chevron nozzle, was applied to the core exhaust of the engine. The experiment in the test cell clarified that the notched nozzle possibly provides better thrust specific fuel consumption than the referential chevron nozzle. The acoustic measurement in the open environment confirmed that the notched nozzle has the noise suppression characteristics expected from previous test results. The perceived noise levels are attenuated by 1.5 dB, which is the same as or better than the referential mixer nozzle.

Central targets for jet engine research activities comprise the evaluation of improved engine components and the assessment of novel engine concepts for enhanced overall engine performance in order to reduce the fuel consumption and emissions of future aircraft. Since CO 2 emissions are directly related to engine fuel burn, a reduction in fuel consumption leads to lower CO 2 emissions. Therefore improvements in engine technologies are still significant and a multi-disciplinary pre-design approach is essential in order to address all requirements and constraints associated with different engine concepts. Furthermore, an increase in effectiveness of the preliminary design process helps reduce the immense costs of the overall engine development. Within the DLR project PEGASUS (Preliminary Gas Turbine Assessment and Sizing) a multi-disciplinary pre-design and assessment competence of the DLR regarding aero engines and gas turbines was established. The application of modern preliminary design methods allows for the construction and evaluation of innovative next generation engine concepts. The purpose of this paper is to present the developed multi-disciplinary pre-design process and its application to three aero engine models. First, a state of the art twin spool mixed flow turbofan engine model is created for validation purposes. The second and third engine models investigated comprise future engine concepts: a Counter Rotating Open Rotor and an Ultra High Bypass Turbofan. The turbofan used for validation is based on publicly available reference data from manufacturing and emission certification. At first the identified interfaces and constraints of the entire pre-design process are presented. An important factor of complexity in this highly iterative procedure is the intricate data flow, as well as the extensive amount of data transferred between all involved disciplines and among the different fidelity levels applied within the smoothly connected design phases. To cope with the inherent complexity data modeling techniques have been applied to explicitly determine the required data structures of those complex systems. The resulting data model characterizing the components of a gas turbine and their relationships in the design process is presented in detail. Based on the established data model the entire engine pre-design process is presented. Starting with the definition of a flight mission scenario and the resulting top level engine requirements thermodynamic engine performance models are developed. By means of these thermodynamic models, a detailed engine component pre-design is conducted. The aerodynamic and structural design of the engine components are executed using a stepwise increase in level of detail and are continuously evaluated in the context of the overall engine system.

The efficacy of integrating the lobed mixer with the core flow deswirling struts to create a single component for reducing the exhaust system length, beyond that attainable through mixer optimization alone, has been investigated. This investigation has been conducted via CFD simulations of a medium-bypass turbofan exhaust system at engine cruise representative conditions. Comparative analysis shows that integration augmented thrust output by about 0.02% while total pressure loss was increased by 3.6%. The aim of the study, to show that this new integrated design would have either minimal impact on or improve exhaust system performance, was confirmed. Comparisons of the flow fields and characteristic quantities downstream of the mixer also showed minimal impact on flow through the nozzle. The deswirling strut was offset by 0.65 D h axially when integrated with the mixer, therefore it can be concluded that the exhaust system ducting could be reduced in length by this same measure — saving engine weight in the process.

Maintenance costs are a substantial contributor to airline operating costs. In this context, understanding, analyzing, and predicting engine performance deterioration is crucial. While diagnostic methods for analyzing the current module and overall engine condition are established in state-of-the-art engine condition monitoring (ECM) systems, deterioration modeling and prognosis are still fields of research. The identification of the contribution of deterioration mechanisms, such as fouling, erosion, and abrasion, to the in-service deterioration poses a key challenge in this area. This paper focuses on a top-down approach for the high pressure compressor (HPC) module. The selected approach is to quantify the contribution of individual deterioration mechanisms to the overall HPC efficiency deterioration based on in-flight measurements. This is accomplished by first using the in-flight measurements to analyze the HPC efficiency loss. Then, the resulting time series of the analyzed HPC efficiency loss are preprocessed. Finally, models of the deterioration mechanisms are fitted to the preprocessed time series. The deterioration models are chosen based on literature references to the respective deterioration mechanisms. As multiple influencing factors affect the deterioration mechanisms, a fleet analysis is conducted to select the model inputs. The fitting process involves a parametric nonlinear regression problem. The outcome is an estimation of the evolution of the deterioration mechanisms over time. This methodology is used to evaluate all available in-service engines of the same type and fleet and to define a fleet model. In the final step, benefits and limitations of the fleet model are investigated.

The blockerless cascade thrust reverser is one of the innovative thrust reverser systems, which replaces the traditionally mechanical blocker door with the aerodynamic blocker door by high-pressure secondary injection, thus significantly reduces the nacelle weight and the complexity of the actuator, and especially suitable for high-bypass-ratio turbofan engine. In order to obtain the optimum performance of a blockerless cascade thrust reverser system and provide the guidance for the design of the blockerless cascade thrust reverser system, a blockerless cascade thrust reverser system was studied in this paper based on the Response Surface Method (RSM), focusing on the effect of different geometric and aerodynamic parameters on the thrust reverser performance. Results show that the secondary injection with high pressure forms the blockage effect to the fan flow, then forces the fan flow to deflect and discharge from the cascade window, realizing the reverse thrust. The thrust reverser performance is mainly affected by fan pressure ratio (FPR), secondary flow pressure ratio (SPR), secondary injection position (X jet ), secondary injection angle (α jet ) and cascade installation angle (β), and the dominated factors are FPR, SPR and X jet . According to the obtained response equation of the thrust reverser performance, the relationship between reverse thrust efficiency and various parameters are clearly described, and performance of thrust reverser can be quickly evaluated. Significant interaction effects exist between different two factors, which must be taken into consideration in the design process of the blockerless cascade thrust reverser system, especially for the interaction effect between FPR and X jet , interaction effect between FPR and β. Optimization design with objective of maximum reverse thrust was carried out to determine the best parameter settings, and reverse thrust ratio η Trev of 60% is achieved under the constraint of the secondary flow ratio.

The MEE (More Electric Engine) is a concept for engine system electrification and is an evolutionary step in engine system design that contributes to the reduction of aviation CO 2 emissions. Mifee (Metering and integrated fuel feeding electrification) and the E3M (Engine Embedded Electric Machine) are the key technologies of the MEE. The purpose of engine thermal management is maintaining the balance between heat generation by the engine system and heat dissipation to the outside of the engine. In recent engine system designs, thermal system design has become an issue because of increased heat generation within the system. For example, a recently developed turbo-fan engine system increases the heat generation by introduction of a fan drive gear system that produces a large amount of heat in addition to the conventional heat source, such as engine main bearings and gears. The MEE will have further heat sources within its system, like the E3M, which is a high-power electric machine. In this paper, an investigation approach and the result of a feasibility study of the MEE thermal management system is described. In addition, the perspective of the technology trend from the MEE toward future hybrid propulsion is also discussed. The global requirements for climate protection strongly demand game-changing technology that significantly improves the aircraft’s overall efficiency. A series/parallel partial hybrid propulsion system, in which both a turbo-fan engine and electrical motor-driven fans generate propulsive power, is considered to be one of the most promising approaches for the future commercial aircraft hybrid propulsion system. The MEE and E3M technology evolves until it will be applied in hybrid propulsion system.

In the present paper, the synergistic combination of inter-cooling with pulsed detonation combustion is analyzed concerning its contribution to NO x and CO 2 emissions. CO 2 is directly proportional to fuel burn and can, therefore, be reduced by improving specific fuel consumption and reducing engine weight and nacelle drag. A model predicting NO x generation per unit of fuel for pulsed detonation combustors, operating with jet-A fuel, is developed and integrated within Chalmers University’s gas turbine simulation tool GESTPAN. The model is constructed using CFD data obtained for different combustor inlet pressure, temperature and equivalence ratio levels. The NO x model supports the quantification of the trade-off between CO 2 and NO x emissions in a 2050 geared turbofan architecture incorporating intercooling and pulsed detonation combustion and operating at pressures and temperatures of interest in gas turbine technology for aero-engine civil applications.

The EU Horizon 2020 ULTIMATE project aims to mitigate one of the greatest loss sources in modern turbofans: the heat in core exhaust gases. The introduction of a closed-circuit recuperated bottoming cycle, using supercritical CO 2 (S-CO 2 ) as a working fluid that is heated by the exhaust gases, has been shown to be a feasible option. Involute spiral heat exchangers have been studied for intercoolers and cooled cooling air systems. However, placing them in the core exhaust and using S-CO 2 implies significant mechanical design challenges from the elevated temperatures and high internal pressures. The studied scenario considers a heat exchanger of 60% effectiveness, with an internal fluid pressure of 32 MPa. The primary objective is to minimise the total mass of the heat exchanger tubes. The work is focused on studying elliptical tubes and an alternative multi-arc cross-section design with internal webs that is more structurally efficient. A parametric analysis of the proposed geometry has been conducted to capture the influence of each of the geometric variables on the resulting stresses. The alloy Ti-6Al-4V is selected as the tube material and the results show that for a 400 MPa maximum allowable stress, a chord of 10 mm, a chord to thickness ratio of eight and a mi nimum wall thickness of 0.2 mm, the minimum tube weight is 20.5 g/m.

With highly integrated airframe architectures emerging as the leading concept of next generation aviation vehicles, research is needed to understand the interactions between inlet swirl distortions and turbofan engines. To meet these research demands, a computational fluid dynamics investigation was conducted to monitor the streamwise development of a complex swirling velocity field in the inlet duct of a turbofan engine with and without the presence of the turbofan nose cone component. By modeling the two geometric setups, natural fluid development and forced fluid/nose cone interactions were distinguishable. To validate the model, computational results were compared to existing experimental data at the fan rotor inlet plane. With the nose cone included, flow angle and swirl intensity predictions from the computational approach agreed well with the experimental measurements. The computational results were expanded upstream to demonstrate the effects of the nose cone geometry on the incoming swirl distortion. Radial flow angles in the presence of the nose cone began to vary from natural swirl development at approximately 0.25 fan diameters upstream, reaching a maximum difference near the leading edge of the nose cone component. Results from this investigation provided a validated model for the prediction of swirl development in a turbofan inlet duct in the presence of a nose cone. Significant change in the swirl profile development was shown from natural vortex motion to induced fluid/solid interactions.

Intercooled recuperated turbofan engines with high bypass ratio are becoming a research focus in recent years due to its advantages of relatively better fuel economy, lower emission and noise characteristic. The re-heater can recover waste heat in the exhaust gas downstream of the low pressure turbine to reduce the specific fuel consumption, and the intercooler can improve compression ability of the compressors with sufficient temperature difference between the high pressure compressor and the low pressure turbine. An optimal pressure ratio split is often sought to maximize the effect of the intercooler on improving the compression ability of the compressors. To determine an optimal pressure ratio split, different combinations of pressure ratio between high and low pressure spools must be calculated, and this requires huge amount of work with the traditional method to achieve the suitable cycle selections. In this paper, theoretic thermodynamic analysis is carried out to derive an explicit solution of the optimum pressure ratio split for maximizing the efficiency of the whole compression path. The effects of different variables on the optimum pressure ratio split are investigated according to the correlated variables in the solution function. A comparison calculation is also made to validate the effectiveness and accuracy of the explicit solution. The results show that the optimum pressure ratio split can be achieved with the derived solution function, which will significantly simplify the process of the cycle parameter selection.

A high bypass ratio turbofan engine capable of powering the Boeing 757 was considered for thrust and drag analysis. A quasi-2D engine model applying the fundamental thermodynamics conservation equations and practical constraints determined engine performance and provided cross-sectional areas in the low-pressure system. Coupled with suggestions on boat-tail angle and curvature from literature, a representative bypass duct and primary exhaust nozzle was created. 3D steady-RANS simulations using Fluent ® 18 were performed on a 1/8 th axisymmetric section of the geometry. A modified 3D fan zone model forcing radial equilibrium was used to model the fan and bypass stator. Takeoff speed and cruise operating conditions were modeled and simulated to identify changes in thrust composition and intake sensitivity. Comparison between net thrust predictions by the engine model and measured in CFD were within grid uncertainty and model sensitivity at cruise. Trends observed in a published database were satisfied and calculations coincided with GasTurb™ 8.0. Verification of thrust in this manner gave confidence to the aerodynamic performance prediction of this modest CFD. Obtaining a baseline bypass design would allow rapid testing of aftermarket components and integration techniques in a realistic flow-field without reliance on proprietary engine data.

The Heron Fan is a new concept of a fuel powered jet engine that does not utilize a conventional core engine. The fan, a single axial compressor of high diameter, creates thrust, similar to a turbofan. Its blades are hollow with inner channels to transport the core air from the hub to the tip, inducing radial compression. The combustion chamber is located in the casing region, either integrated in the blades or in an external ring. After burning, the core air is returned to the blades and is blown out through an expansion device with a large component in circumferential direction. This propels the fan in the opposite direction. The expansion device may be realized by nozzles integrated in the blade trailing edge or by turbine stages integrated in the blade tip region. Subsequently, the core air mixes with the bypass air, which passes the fan axially, and ejects through the main nozzle, producing thrust. To achieve higher compression ratios, it is possible to install core air compressor stages ahead of the fan. The main purpose of this concept is to reduce weight and complexity of the engine, leading to lower production and operating costs. This is achieved by simplifying the engine architecture, integrating the functions and shortening some of the components. In particular, the core engine has been rearranged, thus eliminating the second and in some cases the third shaft. Further, the complete expansion and parts of the compression have been integrated in the fan blade. To assess the aero-thermodynamic parameters, a preliminary cycle analysis has been done, where the most influential parameters were varied. The results show, that the above listed benefits can be achieved while maintaining an efficiency comparable to conventional turbofans. Further, a feasibility study in terms of geometry, internal flow, component implementation and installation has been done, in order to qualify the concept and to identify the most critical aspects. To incorporate the corresponding thoughts and results, as well as to find and eliminate conceptual conflicts and opposing trends, a CAD model has been generated. Overall, the results are sound and encouraging, hence justifying future investigations. However, the Heron Fan concept also brings structural, thermal and aerodynamic challenges which are illustrated and briefly discussed, but still need detailed investigation.

One of the perspective schemes of air breathing engine is a scheme with Ultra High Bypass Ratio (BPR 16...25) Counter Rotating Fan. This solution potentially allows significant increase of fuel efficiency compared to modern conventional turbofans. The model UHBR counter rotating fan named COBRA-1 was developed by CIAM within the framework of European Project COBRA (Innovative Counter rOtating fan system for high Bypass Ratio Aircraft engine). The fan was designed using up-to-date 1D, 2D and 3D methods. COBRA-1 is a 0.7 m diameter model of counter rotating fan driven by a planetary reduction gearbox. The bypass ratio of COBRA-1 is 20. The R2/R1 torque ratio was chosen to obtain 1.42-muliple prevalence in power for 2nd row. The blade numbers are 8/12 for R1/R2 correspondingly. Final geometry of airfoils was defined by 3D profiling process to achieve required aerodynamics and acoustic parameters. Application of control-diffusion airfoils allows reaching high integral performances: specific mass flow equals 211 kg/(s*m^2) and isentropic efficiency at design point is higher than 0.93. The paper presents results of computational simulation of the flow in UHBR fan COBRA-1 based on 3D steady RANS method, 3D URANS and Non-Linear Harmonic method for different operation conditions in comparison with experimental data. Numerical simulation was carried out using Numeca FINE TURBO software package. Steady RANS approach was used during design process to make quick estimation of performances at different rpm. 3D URANS simulation was conducted to analyze unsteady wake-blade and shock-wave interaction and to make a decision about sufficient value of axial gap between rotors. The COBRA-1 fan was tested in CIAM at C3-A test facility which allows conducting a wide range of measurements of local and integral parameters including acoustics of ducted counter rotating fan at different operating conditions. Experimental results demonstrate a high level of integral performances and good agreement with computed values. Significant part of numerical and experimental investigation is devoted to effect of gear-box requirements on aerodynamics. C3-A rig allows to set rotational speed of rotors independently and measure torques at each shaft to achieve required torque ratio and study the influence of small (3–5%) deviation in rpm on aerodynamic characteristics.

There is an increasing dependence on conventional fuels for aviation. In order for a country’s air force to sustain a steady and a secure supply of fuel for aircraft with foresight into the future, alternate sources of fuels must be considered. This paper describes a thermodynamic performance simulation method and analysis of a model military aero gas turbine engine operating in several off-design modes while employing various types of blended fuels between Jet A and alternate bio-fuels, including Synthetic Paraffinic Kerosene (SPK) from Algae (Bio-Algae), Jatropha (JSPK) and Camelina (CSPK). These fuels are already approved by American Society for Testing and Materials (AS™) for blending with 50% Jet A fuel. A thermodynamic performance model for the model engine similar to GE F404-400 turbofan engine has been set up using Pythia, a Cranfield University created performance simulation software implemented with multiple fuel capabilities. The simulated performance and a comparative study shows that the performance of the model engine using blended fuels between Jet A and a bio-fuel were found to be equal or better than that using pure conventional fuel Jet A. The key hot section temperatures, pressures and the fuel consumption of the model engine were found to be slightly lower with the blended fuels than that using Jet A.