Low Bypass Ratio Turbofan Engine Research Articles

Integrated Design of a Variable Cycle Engine and Aircraft Thermal Management System

Abstract The integrated design of a variable cycle engine (VCE) and an aircraft thermal management system (TMS) is investigated. The integrated system is designed using the multiple design point approach in order to achieve required performance metrics at points other than the cycle design condition. The VCE architecture is a three stream design where the third stream is split off after the fan, exhausting through a separate third-stream nozzle. The primary air stream passes through a low-pressure compressor before splitting into an inner bypass stream and a core stream. The inner streams mix aft of the low-pressure turbine and exhaust through a core nozzle. The variable cycle engine utilizes variable compressor inlet guide vanes, a variable area bypass injector at the core stream mixing plane, and variable throats in the two exhaust nozzles. The TMS architecture is an air cycle system using air bled from the high-pressure compressor. The effect of integrating the TMS into the engine design loop is investigated. A comparison is made to prior studies where the same TMS architecture was connected to a low bypass ratio turbofan engine. The comparison shows that the variable cycle engine is able to improve heat dissipation capability versus a ram air cooled system, while eliminating the airframe integration impact that comes with a separate ram-air stream. Lastly, the impact of modulating the variable geometry features on overall cooling capability is investigated. Results are presented for individual operating points as well as at the aircraft mission level.

Is an elliptic jet quieter than a round jet?

Non-axisymmetric geometries, mainly elliptic and rectangular, have been proposed for the reduction of jet noise vis-à-vis round nozzles. Most of the studies of these nozzles are from unheated jets and are restricted to nozzles of very small size. Furthermore, all of them have been carried out at static conditions, thereby rendering their value to insignificance for practical applications. All engines in service with long ducts and a confluent nozzle incorporate an internal lobed mixer. The aeroacoustic characteristics of an elliptic compound nozzle that represents the geometry of an existing low bypass ratio (BPR) turbofan engine, is investigated at 1/7th scale in this study. Typical engine cycle conditions are chosen; data are acquired statically and in the presence of a flight stream. The aspect ratio of the nozzle is 2.0; higher aspect ratios are not suitable for engine applications. The results are compared with a round compound nozzle with the same internal geometry, so as to assess the acoustic benefit, if any, of the elliptic nozzle. Both a simple internal splitter and an in-service lobed mixer have been considered. The elliptic nozzle introduces azimuthal asymmetry even for an unheated jet; the magnitude of azimuthal variation becomes pronounced for heated jets. Typically, the lowest level of noise is observed towards the narrow side of the elliptic nozzle (ϕ = 0°); the noise level gradually increases and reaches a maximum towards the broader side (ϕ = 90°). Though there are some superficial similarities between the elliptic and beveled nozzles, it is shown that the noise characteristics are very different. A systematic study is carried out, with step-by-step build up to realistic geometry, with forward flight. A large noise reduction of ∼3 to ∼4 EPNdB is observed for the splitter nozzle under static conditions. The introduction of a realistic lobed mixer reduces this benefit to close to zero. Finally, there is a noise increase at all azimuthal angles with forward flight. Therefore, the elliptic nozzle does not provide any EPNL benefit for actual nozzle geometry and consequently does not constitute a viable design for noise reduction. The importance of evaluating noise reduction concepts using appropriate geometry and under realistic forward flight conditions is emphasized once again.

Optimal Sparse Volterra Modeling for Transient Behavior of Turbofan Engines based on Internet of Things

In recent years, Internet of Things (IoT) technologies have been increasingly utilized to collect enormous volumes of performance data for intelligent analysis and modeling of aero-engines. This has aided in the development of numerous data-driven solutions so that a strong knowledge of the intricate operations within the equipment is no longer needed. To characterize the dynamic nonlinear transient behavior of turbofan engines, fast response changes have the possibility of being captured accurately through high-order, long-memory-length Volterra series. However, exponentially increasing coefficients are still challenging to be handled properly. For fast and reliable modeling of turbofan engines, an Optimal Sparse Volterra (OSV) model is developed in this paper by reconstructing sparse nonzero coefficients after a global selection through particle swarm optimization. The OSV model focuses on the optimal sparsity of the Volterra kernels while being insensitive to the signal length. Besides, noise reduction and the correlation analysis method are specifically designed for sensor measurements of low-bypass ratio turbofan engines. The OSV model, as well as retaining the powerful descriptive capability of the Volterra series for nonlinear characteristics, finds the most relevant sets of variables and the set of model parameters automatically under the minimum computing workload. According to the experimental results, when real test data are used for turbofan transient maneuvers, the OSV model ensures that the mean absolute error is less than [Formula: see text] for high-pressure rotor speed, thrust and exhaust temperature. Moreover, the nonzero identification coefficients produced by the OSV model in the experiments are less than 6% of the total coefficients. At the same time, the average running time required by the OSV model is less than 35% of that of traditional identification algorithms.

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.

Performance Degradation Evaluation of Low Bypass Ratio Turbofan Engine Based on Flight Data

A low bypass ratio turbofan engine operates in a hostile environment, resulting in performance degradation. This seriously affects the security and reliability of the engine. Therefore, a performance degradation evaluation method for engines based on flight data is proposed. The method expands the equation system to solve the underdetermined problem caused by the lack of engine sensors based on multiple operating point analysis. The improved evolution algorithm is employed to solve the equation system, which relieves the problem of insufficient precision. The engine performance degradation dataset is established based on the engine performance calculation model to verify the reliability of the degradation evaluation method. The results show that the method is applicable to the dataset. Finally, the method is applied to the actual flight data to study the law of the performance degradation of the researched engine, which indicates that the engine’s fan efficiency and high-pressure compressor flow capacity have an apparent downward trend over time.

Design and Analysis of an Aircraft Thermal Management System Linked to a Low Bypass Ratio Turbofan Engine

Abstract The design of an aircraft thermal management system (TMS) that is capable of rejecting heat loads into the bypass stream of a typical low-bypass ratio turbofan engine, or a ram-air stream, is investigated. The TMS consists of an air cycle system, similar to the typical air cycle machines used on current aircraft, both military and commercial. This system turbocharges compressor bleed air and uses heat exchangers in a ram air stream, or the engine bypass stream, to cool the engine bleed air prior to expanding it to low temperatures suitable for heat rejection. In this study, a simple low-bypass ratio afterburning turbofan engine was modeled in numerical propulsion system simulation to provide boundary conditions to the TMS system throughout the flight envelope of a typical military fighter aircraft. Two variations of the TMS system, a ram air cooled and a bypass air cooled, were sized to handle a given demanded aircraft heat load. The ability of the sized TMS to reject the demanded aircraft load throughout several key off-design points was analyzed. It was observed that the maximum load dissipation capability of the TMS is tied to the amount of engine bleed flow, while the level of bleed flow required is set by the temperature conditions imposed by the aircraft cooling system. Notably, engine bypass stream temperatures significantly limit the thermodynamic viability of a TMS designed with bypass air as the heat sink. The results demonstrate the advantage that variable cycle engines (VCEs) may have for future aircraft designs.

The effect of inlet distortion on low bypass ratio turbofan engines

Inlet flow non-uniformity, commonly known as inflow distortion, has been a long-standing problem in the history of gas turbine engines. Distortion can be present in the form of total pressure, total temperature or inflow incidence or any combinations of these. The search for better and robust performance requires engines that can sustain a large amount of inlet distortion without considerable loss in the thrust. In the present paper, the effect of total pressure distortion on a single-stage compressor and low bypass ratio fans are studied. Distortion near hub and tip in the form of step radial total pressure profiles is imposed at far upstream of the rotor leading edge. A systematic approach to qualitatively predict the performance maps in the presence of these distortions is discussed. Further, two extents of total pressure distortion are explored for constant inlet distortion intensity. Hub distortion is found to increase the stability margin, whereas tip distortion reduces it. On extending the distortion extent, hub distortion drastically reduces the stability margin, whereas a comparatively lower reduction in stability margin with tip distortion is observed. The critical distortion limit is observed by varying the inlet distortion extent. Also, it is found that downstream ducts in the bypass axial fan do not interact with the upstream fan. This can be exploited to perform independent simulations of the core engine from low bypass ratio fans. Hub distortion is found to drastically affect the duct performance owing to the presence of thicker upstream inlet boundary layer.

Adaptation Method for Overall and Local Performances of Gas Turbine Engine Model

An adaptation method was proposed to improve the modeling accuracy of overall and local performances of gas turbine engine. The adaptation method was divided into two steps. First, the overall performance parameters such as engine thrust, thermal efficiency, and pressure ratio were adapted by calibrating compressor maps, and second, the local performance parameters such as temperature of component intersection and shaft speed were adjusted by additional adaptation factors. An optimization technique was used to find the correlation equation of adaptation factors for compressor performance maps. The multi-island genetic algorithm (MIGA) was employed in the present optimization. The correlations of local adaptation factors were generated based on the difference between the first adapted engine model and performance test data. The proposed adaptation method applied to a low-bypass ratio turbofan engine of 12,000 lb thrust. The gas turbine engine model was generated and validated based on the performance test data in the sea-level static condition. In flight condition at 20,000 ft and 0.9 Mach number, the result of adapted engine model showed improved prediction in engine thrust (overall performance parameter) by reducing the difference from 14.5 to 3.3%. Moreover, there was further improvement in the comparison of low-pressure turbine exit temperature (local performance parameter) as the difference is reduced from 3.2 to 0.4%.

Combining effect of optimized axial compressor variable guide vanes and bleed air on the thermodynamic performance of aircraft engine system

Aim of this work is to provide evidence of the effectiveness of combined use of the variable guide vanes (VGVs) and bleed air on the thermodynamic performance of aircraft engine system. This paper performed the comparative study to evaluate the overall thermal performance of an aircraft engine with optimized VGVs and bleed air, separately or simultaneously. The low-bypass ratio turbofan engine has been modeled with a 0D/1D modeling approach. The genetic algorithm is employed to find the optimum schedule of VGVs and bleed air. There are four types of design variables: (1) the inlet guide vane (IGV) angle, (2) the IGV and 1st stator vane (SV) angles, (3) bleed air mass flow rate at the exit of the axial compressor, and (4) both type 2 and type 3. The optimization is conducted with surge margin constraints of more than 10% and 15% in the axial compressor. The results show that the additional use of the bleed air increases the efficiency of the compressors. Overall, the percentage reductions of the total fuel consumption for the engine with the IGV, 1st SV and bleed air schedule is 1.63% for 15% surge margin constraints when compared with the engine with the IGV schedule.

A full engine cycle analysis of a turbofan engine for optimum scheduling of variable guide vanes

This paper proposes a full engine cycle analysis to derive the schedule of variable guide vanes (VGVs) in a multi-stage axial compressor for improving the performance of a turbofan engine with required surge margin. Most researchers assumed the operating line of a compressor in process of scheduling of VGVs. However, it has been known that the performance of a compressor is influenced by the angle of VGVs. As a result, the performance variation of a compressor can affect an engine. Therefore, a full engine cycle analysis is conducted to consider the effect of VGVs on a turbofan engine. The VGVs are applied to the front stages of the high pressure compressor in the low-bypass ratio turbofan engine. The compressor of the front stages is transonic three-stage axial flow compressor. The commercial engine simulation program, NPSSTM is employed to analyze the on- and off-design performance of the turbofan engine. 1D meanline analysis is used for performance prediction of a multi-stage axial flow compressor with VGVs. The engine simulation program is coupled with the compressor performance prediction tool to consider the engine performance variation with application of the VGVs. The proposed scheduling algorithm of the inlet guide vane (IGV) and stator vanes (SV) is adjusting the vanes angle to have the minimum loss incidence angle at all rotor blades and to satisfy the required surge margin at off-design condition. The result shows that the proposed algorithm can obtain the scheduling of the IGV, 1st and 2nd SVs angle with stable operation and improved specific fuel consumption of the engine.

Noise reduction in supersonic jets by nozzle fluidic inserts

Professor Philip Doak spent a very productive time as a consultant to the Lockheed-Georgia Company in the early 1970s. The focus of the overall research project was the prediction and reduction of noise from supersonic jets. Now, 40 years on, the present paper describes an innovative methodology and device for the reduction of supersonic jet noise. The goal is the development of a practical active noise reduction technique for low bypass ratio turbofan engines. This method introduces fluidic inserts installed in the divergent wall of a CD nozzle to replace hard-wall corrugation seals, which have been demonstrated to be effective by Seiner (2005) [1]. By altering the configuration and operating conditions of the fluidic inserts, active noise reduction for both mixing and shock noise has been obtained. Substantial noise reductions have been achieved for mixing noise in the maximum noise emission direction and in the forward arc for broadband shock-associated noise. To achieve these reductions (on the order of greater than 4 and 2dB for the two main components respectively), practically achievable levels of injection mass flow rates have been used. The total injected mass flow rates are less than 4% of the core mass flow rate and the effective operating injection pressure ratio has been maintained at or below the same level as the nozzle pressure ratio of the core flow.

Adaptation of the Beveled Nozzle for High-Speed Jet Noise Reduction

The noise of high-speed jets, generated by nozzle exhaust systems that resemble typical jet engine installations on fighter aircraft, has been investigated with the main objective of developing noise reduction designs. To this end, the aeroacoustic characteristics of beveled nozzles have been examined in model scale tests and their performance relative to a conventional round nozzle system has been established. Two different physical models with appropriate area ratios to simulate the engine operating conditions at different power settings were tested; these correspond to 96% N1 (Military or MIL power) and 91% N1 (cutback power). Aeroacoustic measurements and analyses have been carried out for nozzles at bevel angles of 20o, 24o, 28o and 35o for MIL power, and 24o and 32o for cutback power. For the same plenum conditions, the mass flow rates for the beveled nozzles are identical to those of the baseline round nozzle, for a wide range of NPR. The thrust coefficients are higher for the beveled nozzles for both the nozzle geometries at typical MIL power and cutback power, respectively. The increase in the thrust coefficient ranges from ~0.75% for the bevel35 to ~2% for bevel28, in spite of the slight deflection of the jet plume. The reasons for this phenomenon are examined. The beveled nozzles produce at least the same or greater absolute thrust as the baseline nozzle. The beveled nozzles specifically reduce the noise in the peak polar radiation angles; the maximum noise benefit is observed in the azimuthal direction of the longer lip. The magnitude of noise reduction increases with increasing bevel angle, with the largest reduction for bevel35. An examination of the dBA metric, which is used in noise exposure studies, indicates that there is a noise benefit of ~3 dBA to ~4 dBA in the peak radiation sector. The noise benefit in the azimuthal direction of the shorter lip is only slightly lower than that in the direction of the longer lip. I. Introduction he noise of high-speed jets, as from the engines that power fighter aircraft, is a major concern for both military personnel and civilians living close to military bases. Military jets are powered by turbojets or very low bypassratio turbofan engines (BPR ≤ 0.3). The attendant high jet velocities produce extremely high levels of jet noise. This problem is severe for operations on aircraft carrier decks because military personnel are stationed very close to the aircraft during the launch and the landing of carrier-based fighter aircraft. Pilots perform training missions, called Field Carrier Landing Practice, at military bases that mimic the actual flights at the appropriate engine power settings, thereby creating a huge noise problem for the surrounding communities. The magnitude of the problem is first highlighted with a comparison of the noise footprints from the F/A-18 A/B and the Boeing 737-800 aircraft. The noise foot print is calculated as follows: first, A-weighted spectra are calculated from the measured spectra over a wide range of radiation angles. The A-weighting adjusts the full-scale spectrum measured by the microphone to account for the response of the human ear, with the low frequency levels decreased by several decibels, and the high-frequency portion of the spectrum relatively unmodified. The energy levels in the resulting spectrum at the different frequencies are logarithmically summed to produce one dBA number at each angle. The maximum dBA

Model-based on-board turbofan thrust estimation

In this article model-based thrust estimation for turbofan engines is discussed. Two different model-based thrust estimation filters are applied to a low bypass ratio turbofan engine. The two estimators are based on a simple gross thrust model and on a thermodynamic semi-transient model, respectively. The filter parameter estimation as well as the corresponding validation of parameters is based on the use of the “leave one out cross validation” technique. In this article it is demonstrated, in particular, that a thrust estimation filter based on the thermodynamic semi-transient engine model 1 1 Steady state model with measured dynamic states as input. shows good accuracy and predicts the transient thrust and static measurements very well.

Community Noise Levels of the DC-10 Aircraft

The results of recent flyover noise tests show that the McDonnell Douglas DC-10 aircraft powered by General Electric CF6-6D engines is much quieter than the current jet transport aircraft powered by turbojet or low-bypassratio turbofan engines. Several major design features incorporated into the engine and the installation of the engine on the aircraft have accomplished this reduction in noise. The DC-10 will meet the noise level requirements established by the Federal Aviation Administration for new transport aircraft and will, in fact, generate noise levels which are well below the requirements for takeoff and sideline noise. It will also meet the noise level requirements established by the Port of New York Authority and other airport operators at specific noise monitoring stations for normal aircraft operations. The DC-10 represents a major step in the direction of reducing noise pollution in communities around airports.