Browse Topic: Turbofan engines

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Performance Analysis of a High-Speed Tiltrotor Aircraft Featuring VerdeGo Aero's VH-5 Hybrid-Electric Turbofan EngineF-0080-2024-12355/7/2024
In the pursuit of High-Speed Vertical Takeoff and Landing (HSVTOL) platforms, VerdeGo Aero offers its VH-5 hybrid-electric turbofan as the answer to missions requiring high power, vertical lift, and jet-like speeds. To evaluate the possibility of designing a real HSVTOL aircraft around VerdeGo's VH-5 powerplant, this paper investigates the size and expected performance of a militarized spinoff of NASA's Class B, High-Efficiency Civil TiltRotor (HECTR) concept, which has been renamed the VerdeGo Hybrid-Electric Combative TiltRotor, or "VHECTR" for short. Through an in-depth conceptual weight buildup of four commonly proposed tiltrotor architectures, this paper suggests that an entirely new, turbofan-driven propulsion system is needed if modern day HSVTOL demands are to be met. Hence, a separate, yet more conventional "Modified HECTR" tiltrotor configuration is considered to contest the proposed, VH-5 powered VHECTR concept. However, the results of a full-scale aircraft comparison, aerodynamic analysis, and mission analysis alongside the V-22 Osprey indicate the VHECTR concept is more HSVTOL mission capable with greater time and fuel savings than any advanced tiltrotor system to date.
Roiati, Riccardo, Rice, Tyler, Steinfeldt, Brock, Nitzsche, Jasper, Anderson, Richard
Landing Gear Integration into Aircraft Structure in Early Design Stage2019-01-18909/16/2019
The demanded development towards various emission reduction goals set up by several institutions forces the aerospace industry to think about new technologies and alternative aircraft configurations. With these alternative aircraft concepts, the landing gear layout is also affected. Turbofan engines with very high bypass ratios could increase the diameter of the nacelles extensively. In this case, mounting the engines above the wing could be a possible arrangement to avoid an exceedingly long landing gear. Thus, the landing gear could be shortened and eventually mounted at the fuselage instead of the wings. Other technologies such as high aspect ratio wings have an influence on the landing gear integration as well. To assess the difference, especially in weight, between the conventional landing gear configuration and alternative layouts a method is developed based on preliminary structural designs of the different aircraft components, namely landing gear, wing and fuselage. Simplified parametric finite element structural models for the different components are introduced. These models are used to investigate different aircraft configurations with special regard on the landing gear integration. The structural models of the fuselage and the wing are sized according to defined load cases. After this first sizing step, the structural model of the landing gear is integrated and different landing gear load cases are applied. The developed methods aim to assess the impact of the different landing gear configurations, not only regarding the landing gear structure but also the surrounding support structure of the airframe. Results of the applied methods for aircraft configurations with different landing gear integration are presented and discussed.
Kling, Ulrich, Hornung, Mirko
Aircraft Engine Icing Event Avoidance and MitigationTBMG-347387/1/2019
Innovators at NASA’s Glenn Research Center have developed a new means of avoiding and mitigating icing events for aircraft flying above 14,000 feet, dramatically improving aviation safety and reducing operating costs.
Wind Tunnel Measurements of Simulated Glaciated Cloud Conditions to Evaluate Newly Developed 2D Imaging Probes2019-01-19816/10/2019
Instrumentation that has been used for characterization of mixed-phase and glaciated conditions in the past, like the OAP probes, are subject to errors caused by variations in diffraction on the images away from the object plane and by the discrete nature of their particle detection and sizing. Correction methods are necessary to consider their measurements adequate for high ice water content (IWC) environments judged to represent a significant safety hazard to propellers and turbofan engine operability and performance. For this reason, within the frame of EU FP7 HAIC project, instrumentation characterization and validation is considered a major element need for successful execution of flight tests campaigns. Clearly, instrumentation must be sufficiently reliable to assess the reproducibility of artificial clouds with high ice water content generated in icing tunnels. Instruments are required to measure these conditions with a sufficient level of accuracy for the purposes of the testing. Currently, there is an anticipated basic uncertainty of a factor of 2-5 when measuring clouds in-situ. This may be worse for thunderstorm core regions, because of the poorly measured ice particle sizes below 100 μm when using legacy instruments. Measurements below 100 μm are especially difficult for ice particles. In order to mitigate this measurement limitation, an innovative approach for imaging droplets and ice crystals was considered in this work. This method is based on Artium Technologies Inc. High Speed Imaging (HSI) instrument. The method utilizes a unique multi-beam illumination approach to control the depth of field, minimize out-of-focus image noise effects, and limit obscuration of particles in the sample volume produced by out-of-focus particles in the optical path. Rather than utilizing a linear array to acquire the images, a modern CMOS two-dimensional array imaging system is used. The lasers used for producing bright field shadow images are pulsed synchronously and have pulse duration of as little as 12.5 ns to “freeze” the particle motion. The response of the HSI will be reviewed on the basis of comparison with SPEC 2D-S by means of wind tunnel measurements for glaciated cloud conditions.
Esposito, Biagio M., Bachalo, William D., Leroy, Delphine, Schwarzenboeck, Alfons, Jurkat, Tina, Voigt, Christiane, Bansmer, Stephan
The Effect of Crosswinds on Icing at Turbofan Engine Inlet2019-01-20246/10/2019
Ice accretion at engine inlet has a dangerous effect on the inlet airflow and shed ice would be ingested into the engine and cause compressor blades damage, and even combustors flame out. In order to analyze the effect of crosswinds on icing at turbofan engine inlet, a complete icing analysis method, which is based on the Messinger model and takes the influence of runback water into consideration, is constructed. The runback water is considered laminar flow and the flow direction is dominated by the bottom flow of air. The supercooled water droplets impingement, ice accretion and runback water characteristics and inlet distortion with and without ice were investigated at crosswinds speed of 15, 20, 25, 30 kt. The variation of local water collection coefficient β is unchanged at crosswind conditions, but the location of the maximum value and non-dimensional impact limits change with the crosswind, which move backward along the outer edge on the windward, and move along the inner edge on the leeward. The ice thickness on windward surface increases slightly near the stagnation point, and the outer icing limit occurs at more rear position with crosswinds. On leeward surface, the variation is opposite to the windward and the ice shapes present a more symmetrical distribution at stagnation point. On the smooth inlet surface, the inlet airflow at the fan is relatively uniform, and the total pressure ratio variation at the inlet lip also is regular as crosswind increases. Greatly losses of airflow occur after ice accretion at crosswind conditions, which cause uneven inlet airflow. The effect of the crosswind on ice accretion at the inlet should be considered when designing and optimizing anti-icing system. Otherwise, the ice formation under crosswind will deteriorate the inlet airflow quality and pose a hazard to aircraft safety.
Yang, Qian, Guo, Zhiqiang, Zheng, Mei, Dong, Wei
Microwave Technique for Liquid Water Detection in Icing Applications2019-01-19306/10/2019
The partial melting of ingested ice crystals can lead to ice accretion in aircraft compressors, but accurately measuring the relatively small fraction of liquid water content in such flows is challenging. Probe-based methods for detecting liquid water content are not suitable for deployment within turbofan engines, and thus alternatives are sought. Recent research has described approaches based on passive microwave sensing. We present here an approach based on active microwave transmission and reflection, employing a vector network analyzer. Utilization of both transmission and reflection provides additional data over and above emission or transmission only, and permits a more controllable environment than passive sensing approaches. The paper specifically addresses the question of whether such an approach is viable within the context of representative icing wind tunnel and engine flow conditions. A quasi-thermal equilibrium approach is presented herein to estimate the melting ratio during microwave analysis of samples at 0 °C. Experimental results using microwaves in the 2.45GHz region are presented, and post-processing methods investigated. This is followed by an investigation of detection limits for ice accretion in the sub-gram range. The results indicate the potential of the technique, with a number of avenues evident for further research.
Leis, John, Buttsworth, David, Saeed, Ramiz, Saleh, Khalid, McGilvray, Matthew, Gillespie, David
Scaling Evaluation of Ice-Crystal Icing on a Modern Turbofan Engine in PSL Using the COMDES-MELT Code2019-01-19206/10/2019
This paper presents preliminary ice-crystal icing (ICI) altitude scaling evaluation results of a Honeywell Uncertified Research Engine (HURE) that was tested in the NASA Glenn Research Center Propulsion Systems Laboratory (PSL) during January of 2018. This engine geometry features a hidden core design to keep the core less exposed. The engine was fitted with internal video cameras to observe various ice buildup processes at multiple selected locations within the engine core flow path covering the fan stator, the splitter-lip/shroud/strut, and the high pressure compressor (HPC) variable inlet guide vane (IGV) regions. The potential ice accretion risk was pre-determined to occur by using NASA’s in-house 1D Engine Icing Risk assessment code, COMDES-MELT. The code was successful in predicting the risk of ice accretion in adiabatic regions like the fan-stator of the HURE at specific engine operating points. However at several operating points during the test, liquid water was observed running along the shroud toward the variable IGV of the HPC regions with an air temperature well below freezing, thus no particle melting could have occurred due to heating from the air alone. It was reasoned that other sources of heat were present in that region. To account for these heat sources the inlet total temperature was adjusted to give a wet bulb temperature of 24 °F below the standard minimum wet bulb temperature of 492 °R to allow ice to accrete in the splitter-lip/shroud/strut region, which was determined from a reference case where hard ice was observed in that region. With that adjustment the COMDES-MELT code was successful in providing operating points where there was a risk of ice accretion during the test campaign. In addition to calculating possible conditions at different selected lower altitudes, simulations were run to determine potential inlet conditions that could lead to ice-crystal accretion along the prescribed stations where the cameras were available. From there, scaled test conditions were determined by best matching the following three icing related parameters of the reference condition: (1) the local air total wet bulb temperature, (2) the local ice crystal cloud melt ratio and (3) the engine fan face ice/water to air mass flux ratio of the ice crystal cloud. Instantaneous images taken from the time-lapsed movies of ice buildup were used along with the relevant thermodynamic data of air, water vapor and local icing condition to help evaluate how closely the proposed altitude scaling method could be used in ground based test facility to duplicate selected reference ICI features observed at specific location inside this engine at different scale altitudes. Discussions on observed limitation for engine icing scaling application from this test campaign and needed improvement are provided. A scaling test procedure to help identify potential ICI risk conditions and possible ice accretion locations of a new turbofan engine is evaluated in PSL.
Tsao, Jen-Ching
Semi-Empirical Modelling of Erosion Phenomena for Ice Crystal Icing Numerical Simulation2019-01-19676/10/2019
The aim of this work is to develop a semi-empirical model for erosion phenomena under ice crystal condition, which is one of the major phenomena for ice crystal accretion. Such a model would be able to calculate the erosion rate caused by impinging ice crystals on accreted ice layer. This model is based on Finnie [1] and Bitter [2] [3] solid/solid collision theory which assumes that metal erosion due to sand impingement is driven by two phenomena: cutting wear and deformation wear. These two phenomena are strongly dependent on the particle density, velocity and shape, as well as on the surface physical properties such as Young modulus, Poisson ratio, surface yield strength and hardness. Moreover, cutting wear is mostly driven by tangential velocity and is more effective for ductile eroded body, whereas deformation wear is driven by normal velocity and is more effective for brittle eroded body. Several researchers based their erosion modelling on these two phenomena such as Hutchings et al. [4] for deformation erosion, or Huang et al. [5] and Arabnejad et al. [6] for cutting and deformation erosion. The main work of this paper is to develop an erosion model for ice crystal impingement based on these two phenomena, and to show its capability to predict accretion shape by simulating experimental cases from the National Research Council of Canada (NRC). NRC’s Currie et al. ice crystal experiments [7] [8] realized in warm aerodynamic conditions, such as the one encountered in high icing severity areas of a turbofan engine, show accretion severity for a large range of liquid water content to total water content. In order to validate the erosion model based on solid/solid collision, this paper presents the simulation of the lower melting rate experiment. Results show fair agreement with experimental data and allow us to propose pertinent further work.
Charton, Virgile, Trontin, Pierre, Aouizerate, Gilles, Villedieu, Philippe
Design of Composite Materials for Improved Loss-of-Lubrication Survivability of Hybrid Rotorcraft GearsF-0075-2019-146855/13/2019
Hybrid gears featuring steel teeth mated to a composite body provide the potential for significant weight savings in aerospace applications such as rotorcraft and geared turbofan engines. For hybrid gears to be viable for use in these applications, they must not degrade mechanical performance or thermal characteristics, particularly under loss-of- lubrication operation. The heat generated by loss-of-lubrication operation may be especially problematic for the fiber- reinforced polymer composite materials used in hybrid steel-composite gear prototypes. Initial hybrid gear design and testing is described in the literature, but no prior studies have investigated optimization of the composite material for thermal performance. In the study presented herein, conductive fibers and high-temperature polymers are introduced to make a composite material better suited to high-temperature and loss-of-lubrication. Channel-flow resin transfer molding was used to fabricate composites with a variety of configurations, including hybrid reinforcement alternating highly-conductive pitch-based carbon fiber plies with typical high-strength polyacrylonitrile (PAN)-based carbon fiber plies. Performance was evaluated by thermal conductivity measurement and compression strength testing. Finite element modeling was performed to quantify strength and stiffness requirements for the composite gear web and to investigate the effects of using plies of different stiffness in a hybrid laminate. Results of this study show that the in-plane thermal conductivity of composite laminates can be greatly increased by using hybrid reinforcement, but the hybrid gear web design must account for the low compression strength of the hybrid reinforcement to ensure that the structural integrity of the composite gear web is not compromised.
Waller, Matthew, Koudela, Kevin, McIntyre, Sean
Novel Turboelectric Aircraft DesignTBMG-288615/1/2018
NASA Langley Research Center has developed a novel aircraft design that can carry 150+ passengers over a range of 3,500 nautical miles. Key features include a turboelectric propulsion system, twin underwing turbofans, and a rear-fuselage, electrically driven, boundary layer ingesting fan. The design is developed for single-aisle-class aircraft (the largest commercial transport class in the world) to significantly reduce fuel burn compared to current aircraft.
Variable Bypass Turbofan EngineTBMG-288775/1/2018
Many aircraft need to operate efficiently at more than one flight regime; for example, certain airframes are expected to perform in relatively high-speed cruise modes, as well as slower loitering scenarios. Unfortunately, an engine operates most efficiently when the exit velocity closely matches the speed of the airframe. As a result, designing an engine that is suitable for multiple roles yields an engine that only performs moderately well in each of those operating conditions.
Software Helps Optimize World’s Largest All-Composite AircraftTBMG-2800812/1/2017
Collier Research Corp. Newport News, VA
ASTM-D7684-11 Compliant Computer Aided Interactive Wear Debris Particle Analysis for On-Site Condition Monitoring, Diagnostics and Prognostics2017-01-21249/19/2017
Within the aviation industry analysis of wear debris particles recovered from magnetic plugs and lubricating fluids is an essential condition monitoring tool. However, in large organisations, high staff turnover in remote work environments often leaves dangerous gaps in on-site support and background knowledge. The current work develops interactive software for wear debris particle classification, root cause diagnosis and serviceability prognostics. During the research several hundred wear debris particle images were collected, analysed and classified by a number of experts. At each stage of the analysis the experts were questioned about the knowledge and experience used to make their diagnoses and prognoses. The end result is an extensive knowledge base representing the combined expertise of a number of highly trained engineers, each with decades of hands-on experience. Access to this knowledge base is provided at each stage of interaction with the software, as automatically generated expert-guidance notes, which support the end user in their decision making by letting them know exactly what the expert would be thinking at that point in the analysis. This helps to correct systematic misconceptions, resolve ambiguities and offers appropriate alternative diagnostic choices. The use of brief, contextual pieces of information means that facts and concepts are more easily understood and remembered. The software uses the nomenclature of, and is compliant with, the ‘ASTM D7684-11 standard guide for microscopic characterisation of particles from in-service lubricants’. It allows the end user to integrate standardised particle analysis into the daily work flow, whilst acquiring knowledge and skills in a practical setting.
Leavers, Violet
Turbofan Engine Acoustic Liner Design and Analysis ToolsTBMG-268665/1/2017
NASA Langley Research Center has developed two tools for turbofan engine acoustic liner design and analysis. The first is a statistical approach for broadband liner design and assessment; the second is graphical software to design and analyze resonant channels in acoustic liners.
Uncertainty of In-Flight Thrust DeterminationAIR1678B (Current)10/22/2016
This document defines and illustrates the process for determination of uncertainty of turbofan and turbojet engine in-flight thrust and other measured in-flight performance parameters. The reasons for requiring this information, as specified in the E-33 Charter, are: determination of high confidence aircraft drag; problem rectification if performance is low; interpolation of measured thrust and aircraft drag over a range of flight conditions by validation and development of high confidence analytical methods; establishment of a baseline for future engine modifications. This document describes systematic and random measurement uncertainties and methods for propagating the uncertainties to the more complicated parameter, in-flight thrust. Methods for combining the uncertainties to obtain given confidence levels are also addressed. Although the primary focus of the document is in-flight thrust, the statistical methods described are applicable to any measurement process. The E-33 Committee has endeavoured to gather industry-wide expertise in in-flight measurement and uncertainty analysis to collect and promulgate recommended practices in the subject disciplines. The Committee is organized into subcommittees to address both the analytical and test methodology for determination of in-flight thrust and also the uncertainty of the determination. This document; Uncertainty of In-flight Thrust Determination, AIR1678, addresses the process for determining the uncertainty of in-flight thrust. A companion document, In-Flight Thrust Determination, AIR1703, addresses the basic methodology for determining in-flight thrust. The Committee, after reviewing recommended changes and clarification in definitions and application of statistical uncertainty items, made small revisions to the original document published in 1985. These changes were incorporated into AIR1678 Rev A. This Revision B has the same Scope as preceding versions. The nomenclature and methodology used herein are now consistent with evolving world and national standards promulgated primarily by ISO and ASME.
E-33 In Flight Propulsion Measurement Committee
Numerical Investigations on a Tunable Helmholtz Resonator: Integration of a Passive Polymer Membrane in a Helmholtz Resonator2016-01-18426/15/2016
A Helmholtz resonator is a passive acoustic resonator used to control a single frequency resulting from the cavity volume and the resonator neck size. The main purpose of work in progress is to propose to investigate numerically some strategies allowing real-time tunability of the Helmholtz resonator in order to provide a wider bandwidth and hence enhance noise attenuation. Two concepts will be developed, both based on the use of electroactive polymer (EAP) membranes. These materials exhibit a change of shape when stimulated by an electric field. The first concept consists in replacing the resonator rigid back plate by an EAP material membrane, while on the second one, the membrane is located in front of the resonator. Numerical investigations are performed using several kinds of a passive EAPs material membranes in order to determine the practical potential of these concepts. Significant enhancements of the transmission loss and/or the absorption properties could be obtained in the low frequency band by tuning the volume and then the resonance frequencies of the resonator. In the current paper, only the passive performances of these two concepts are discussed.
Abbad, Ahmed
Cantera Integration with T-MATSTBMG-240993/1/2016
The Toolbox for the Modeling and Analysis of Thermodynamic Systems (TMATS) software package is a library of building blocks that can be assembled to represent any thermodynamic system in the Simulink® (MathWorks, Inc.) environment. These elements, along with a Newton Raphson solver (also provided as part of the T-MATS package), enable users to create models of a wide variety of systems. The updated version of T-MATS (v1.1.1) includes the integration of Cantera, an open source thermodynamic simulation tool. T-MATS was initially described in detail in LEW-19165-1, "", Software Tech Briefs (September 2014), p. 11.
Automated ATM System Enabling 4DT-Based Operations2015-01-25399/15/2015
As part of the current initiatives aimed at enhancing safety, efficiency and environmental sustainability of aviation, a significant improvement in the efficiency of aircraft operations is currently pursued. Innovative Communication, Navigation, Surveillance and Air Traffic Management (CNS/ATM) technologies and operational concepts are being developed to achieve the ambitious goals for efficiency and environmental sustainability set by national and international aviation organizations. These technological and operational innovations will be ultimately enabled by the introduction of novel CNS/ATM and Avionics (CNS+A) systems, featuring higher levels of automation. A core feature of such systems consists in the real-time multi-objective optimization of flight trajectories, incorporating all the operational, economic and environmental aspects of the aircraft mission. This article describes the conceptual design of an innovative ground-based Air Traffic Management (ATM) system featuring automated 4-Dimensional Trajectory (4DT) functionalities. The 4DT planning capability is based on the multi-objective optimization of 4DT intents. After summarizing the concept of operations, the top-level system architecture and the key 4DT optimization modules, we discuss the segmentation algorithm to obtain flyable and concisely described 4DT. Simulation case studies in representative scenarios show that the adopted algorithms generate solutions consistently within the timeframe of online tactical rerouting tasks, meeting the set design requirements.
Gardi, Alessandro, Sabatini, Roberto, Ramasamy, Subramanian, Marino, Matthew, Kistan, Trevor
Characterization of the Ultrafine and Black Carbon Emissions from Different Aviation Alternative Fuels2015-01-25629/15/2015
This study reports gaseous and particle (ultrafine and black carbon (BC)) emissions from a turbofan engine core on standard Jet A-1 and three alternative fuels, including 100% hydrothermolysis synthetic kerosene with aromatics (CH-SKA), 50% Hydro-processed Esters and Fatty Acid paraffinic kerosene (HEFA-SPK), and 100% Fischer Tropsch (FT-SPK). Gaseous emissions from this engine for various fuels were similar but significant differences in particle emissions were observed. During the idle condition, it was observed that the non-refractory mass fraction in the emitted particles were higher than during higher engine load condition. This observation is consistent for all test fuels. The 100% CH-SKA fuel was found to have noticeable reductions in BC emissions when compared to Jet A-1 by 28-38% by different BC instruments (and 7% in refractory particle number (PN) emissions) at take-off condition. BC emissions from this fuel were lower than from Jet A-1 by 45-50% (and 25-26% in refractory PN) at idle or cruise condition. The 100% CH-SKA fuel was observed to have a minimum influence on non-refractory PN emissions. A lower volume in naphthalene in the 100% CH-SKA fuel was hypothesized to be one of the factors attributing to the reduced BC emissions when compared to Jet A-1 emissions. For the 50% HEFA-SPK fuel, BC emissions were lower than the BC emissions from Jet A-1 by 58-86% for various engine load conditions. BC emissions from the 100% FT-SPK fuel were lower than from the Jet A-1 by 70-98%. Both the refractory and non-refractory PN emissions from these fuels were lower by comparable magnitude when compared to that from Jet A-1.
Chan, Tak W., Chishty, Wajid, Davison, Craig, Buote, David
Computational Method for Ice Crystal Trajectories in a Turbofan Compressor2015-01-21396/15/2015
In this study the characteristics of ice crystals on their trajectory in a single stage of a turbofan engine compressor are determined. The particle trajectories are calculated with a Lagrangian method employing a classical fourth-order Runge-Kutta time integration scheme. The air flow field is provided as input and is a steady flow field solution governed by the Euler equations. The single compressor stage is represented using a cascaded grid. The grid consists of three parts of which the first and the last part are stator parts and the centre part is a rotor. Each particle is modelled as a non-rotating rigid sphere. The remaining model does allow the exchange of heat and mass to and from the particle resulting in a mass, temperature and phase change of the particle. The phase change is based on a perfectly concentric ice core-water film model and it is assumed that the particle is at uniform temperature. The results for the collection efficiency, particle temperature and amount of evaporated mass will be shown for two extreme scenario's. The first simulation is carried out at standard conditions for a Boeing-747 at cruising conditions using the International Standard Atmosphere (ISA) at that altitude, i.e. at 10,650 m. The second simulation is carried out at lower altitude where the existence of supercooled liquid water is thought to be unlikely. Both simulations are carried out at two different temperatures and for either dry or saturated air. The range of particle diameters is set from 10 to 500 micrometres.
Grift, E.J., Norde, E., Van der Weide, E.T.A., Hoeijmakers, H.W.M.
Performance Modeling of Honeywell Turbofan Engine Tested with Ice Crystal Ingestion in the NASA Propulsion System Laboratory2015-01-21336/15/2015
The Propulsion Systems Laboratory (PSL) altitude test facility at NASA Glenn Research Center, has been used to test a full scale Honeywell turbofan engine at simulated altitude operating conditions. The PSL has spray bars to create a continuous cloud of fully glaciated ice crystals. The tests successfully duplicated the icing events that were experienced by the Honeywell engine (ALF502R-5) during flight through ice crystal clouds. After the ice cloud was turned on key engine performance parameters such as the fan speed, air flow rate, fuel flow rate, and compressor exit pressure and temperature responded immediately to the ingestion of the ice crystals. For some of the test points, these performance parameters remained unchanged from the initial response to the ice crystals, while during other test points the engine performance began to deteriorate to the point where an uncommanded loss of thrust control (engine rollback) was judged by the test engineers to have been imminent. The data points where the performance deteriorated with time have been attributed to ice accretion in the low pressure compressor. To gain understanding of the transient engine performance, nine data points were analyzed with the Numerical Propulsion System Simulation engine system code and a compressor flow analysis code. The focus was to improve understanding of the effect of ice crystal ingestion on the transient engine performance, and to quantify the rate of growth of blockage and additional losses due to ice accretion. The engine responds immediately after the initiation of the ice cloud, but whether ice growth in the low pressure compressor will occur after a period of time, appears to be strongly dependent on the concentration of ice crystals ingested, the wet bulb temperature distribution, and the melt ratio.
Veres, Joseph P., Jones, Scott M., Jorgenson, Philip C. E.
Ice Crystal Ingestion in a Turbofan Engine2015-01-21466/15/2015
A through-flow based Monte Carlo particle trajectory simulation is used to calculate the ice crystal paths in the low pressure compressor of a high bypass ratio turbofan engine. The simulation includes a statistical ice particle breakup model due to impact on the engine surfaces. Stage-by-stage ice water content, particle size and particle velocity distributions are generated at multiple flight conditions and engine power conditions. The majority of the ice particle breakup occurs in the fan and first LPC stage. The local ice water content (IWC) within LPC is much higher than the ambient conditions due to scoop effects, centrifuging and flow-path curvature. Also the ice particles approach the stators at lower incidence angles than the air flow. The simulation results prompt the need to revisit the approach for properly setting up boundary conditions for component or cascade testing.
Feulner, Matthew, Liao, Shengfang, Rose, Becky, Liu, Xuejun
Development of a Coupled Air and Particle Thermal Model for Engine Icing Test Facilities2015-01-21556/15/2015
This paper describes a numerical model that simulates the thermal interaction between ice particles, water droplets, and the flowing air applicable during icing wind tunnel tests where there is significant phase-change of the cloud. It has been previously observed that test conditions, most notably temperature and humidity, change when the icing cloud is activated. It is hypothesized that the ice particles and water droplets thermally interact with the flowing air causing the air temperature and humidity to change by the time it reaches the test section. Unlike previous models where the air and particles are uncoupled, this model attempts to explain the observed changes in test conditions by coupling the conservation of mass and energy equations. The model is compared to measurements taken during wind tunnel tests simulating ice-crystal and mixed-phase icing that relate to ice accretions within turbofan engines. The model simulates trends that were experimentally observed, but does not fully explain the measured values. Some possible explanations for this discrepancy are offered. This model, written in MATLAB, is based on fundamental conservation laws and empirical correlations.
Bartkus, Tadas P., Struk, Peter, Tsao, Jen-Ching
Ice Crystal Icing Engine Testing in the NASA Glenn Research Center's Propulsion Systems Laboratory: Altitude Investigation2015-01-21566/15/2015
The National Aeronautics and Space Administration (NASA) conducted a full scale ice crystal icing turbofan engine test using an obsolete Allied Signal ALF502-R5 engine in the Propulsion Systems Laboratory (PSL) at NASA Glenn Research Center. The test article used was the exact engine that experienced a loss of power event after the ingestion of ice crystals while operating at high altitude during a 1997 Honeywell flight test campaign investigating the turbofan engine ice crystal icing phenomena. The test plan included test points conducted at the known flight test campaign field event pressure altitude and at various pressure altitudes ranging from low to high throughout the engine operating envelope. The test article experienced a loss of power event at each of the altitudes tested. For each pressure altitude test point conducted the ambient static temperature was predicted using a NASA engine icing risk computer model for the given ambient static pressure while maintaining the engine speed.
Oliver, Michael
Tool for Turbine Engine Closed-loop Transient Analysis (TTECTrA)TBMG-222056/1/2015
The current steady-state performance-based system studies are not adequate to evaluate certain advanced technologies for the challenges associated with meeting the transient performance requirements without better definition and control of component operating margins. Additional dynamic analyses would require more detailed, physics-based, component-level models and controllers in order to better identify dynamic issues that may arise during operation of the engine. The additional complexity in these models increases the effort necessary throughout the design processes, including the system-analysis step. The solution would be a physics-based engine model with full envelope controller. C-MAPSS and C-MAPSS40k are examples of engine models with full envelope controllers, but because they are designed to model a specific engine, they are not good candidates for systems analysis.
Passive Destructive Interference Acoustic Liner for a Turbofan Engine Using Additive ManufacturingTBMG-219995/1/2015
This technology exploits the capabilities of additive manufacturing to attenuate the fan noise within the inlet or aft duct of a turbofan engine. The approach may be expanded to include auxiliary power units, environmental control systems, or other cooling systems requiring noise attenuation.
Transient Engine Emulation within a Laboratory Testbed for Aircraft Power Systems2014-01-21709/16/2014
This paper presents the details of an engine emulation system utilized within a Hardware-in-the-Loop (HIL) test environment for aircraft power systems. The paper focuses on the software and hardware interfaces that enable the coupling of the engine model and the generator hardware. In particular, the rotor dynamics model that provides the critical link between the modeled dynamics of the engine and the measured dynamics of the generator is described in detail. Careful consideration for the measured torque is included since the measurement contains inertial effects as well as torsional resonances. In addition, the rotor model is equipped with the ability to apply power and speed scaling between the engine and generator. This scaling approach provides significant flexibility that can be useful when hardware resources are limited such that a direct engine-generator match is not possible or when one wants to evaluate turboshaft engine dynamics for a variety of applications and power levels. Depending on what type of scaling is used, there may be an effect on the maximum power the drive stand must provide to fully emulate engine transients. Explicit torque/power requirements are given, along with an evaluation of the existing drive stand's emulation capabilities.
Bash, Michelle, Boyd, Michael, Miller, Chad
HTF7000 Engine Design, Development and Uses2013-01-22289/17/2013
Honeywell has developed a unique turbofan engine for application to the super mid-size business aviation market, the HTF7000. This paper will describe the design of this engine including aeromechanical design of its components. The unique design features of this engine will be described along with the technology growth path to keep the engine current. This paper will also describe several features which have been developed for this engine in response to new regulatory requirements. Some aspects of the engine to aircraft integration will also be described.
Winstanley, David K.
Acoustic Liner for Turbomachinery ApplicationsTBMG-84919/1/2010
The purpose of this innovation is to reduce aircraft noise in the communities surrounding airports by significantly attenuating the noise generated by the turbomachinery, and enhancing safety by providing a containment barrier for a blade failure. Acoustic liners are used in today’s turbofan engines to reduce noise. The amount of noise reduction from an acoustic liner is a function of the treatment area, the liner design, and the material properties, and limited by the constraints of the nacelle or casement design. It is desirable to increase the effective area of the acoustic treatment to increase noise suppression. Modern turbofan engines use “wide-chord” rotor blades, which means there is considerable treatment area available over the rotor tip.
Taking measures of the machineAEROAPR08_034/1/2008
New requirements and technologies advance the art of test and sensing. Few sensor and instrumentation environments are more challenging than those found inside a jet engine. With temperatures up to 1200°F in the hot section, the climate is particularly hostile to just about anything that an engineer can design. “Measurements inside an engine are in a number of ways limited by the operational requirements of the engine,” said Gary Hunter, Intelligence Systems Hardware Lead in the Sensors and Electronics branch at NASA Glenn Research Center (GRC). “It is hot and gases flow at high velocity rates. If you wish something to be in the engine, it needs to survive the harsh environment, and not a lot of things can survive there.”
Rosenberg, Barry
Long Range Aircraft, an Ideal Case for Fuel Cell Power2007-01-295711/28/2007
This paper shows that the analysis of factors such as costs, weights, fuel distribution logistics, etc. indicates that the replacement of internal combustion engines (ICE) by fuel cell/electric motor (FC/EM) systems, can be more attractive to aeronautical than to automotive applications. For long range airplanes especially, where the fuel represents an important fraction of the vehicle weight, the penalties from that replacement by a heavier system, can be overcome by the fuel weight reductions resulting from the higher fuel cell efficiencies. The necessity of FC/EM powered airplanes is also discussed in view of the eventual climate effects the air traffic may have, due to the injection of carbon dioxide and water vapor into the high troposphere. Some of the weight and aerodynamic issues posed by a FC/EM medium sized long range transport airplane preliminary design are discussed, and a non conventional aircraft configuration is proposed to be used in order to partially compensate the increase in the basic operational weight resultant of the FC/EM system adoption.
Galvão, Francisco Leme
System Would Detect Foreign-Object Damage in Turbofan EngineTBMG-10601/1/2006
A proposed data-fusion system, to be implemented mostly in software, would further process the digitized and preprocessed outputs of sensors in a turbofan engine to detect foreign-object damage (FOD) [more precisely, damage caused by impingement of such foreign objects as birds, pieces of ice, and runway debris]. The proposed system could help a flight crew to decide what, if any, response is necessary to complete a flight safely, and could aid mechanics in deciding what post-flight maintenance action might be needed.
Aerospace Engineering 2004-07-01AEROJUL047/1/2004
More-electric small turbofan engines The aerospace industry demands continuous improvement of propulsion engines as well as timely response to the new and changing requirements for them. Zero-emissions aircraft Researchers explore the potential application of alternative fuels and propulsion systems for aircraft that include a business jet and single-engine propeller aircraft. Integration releases the reins Cessna's new Citation Mustang launches the company into a new market for business jets, helped by the latest PLM solutions from IBM. Military: general A military flight-control system proposed for use in general aviation uses inverse control to decouple the flight controls and to modify the handling qualities of the aircraft, while employing articficial neural networks to compensate for any modeling errors.
Program Predicts Broadband Noise From a Turbofan EngineTBMG-17521/1/2004
Broadband Fan Noise Prediction System (BFaNS) is a computer program that, as its name indicates, predicts the broadband noise generated by the fan stage of a turbofan engine. This noise is the sum of (1) turbulent- inflow noise, which is caused by turbulence impinging on leading edges of the fan and the fan exit guide vane and (2) self noise, which is caused by turbulence convecting past the corresponding trailing edges. The user provides input data on the fan-blade, vane, and flow-path geometries and on the mean and turbulent components of the flow field. BFaNS then calculates the turbulent-inflow noise by use of D. B. Hanson's theory, which relates sound power to the inflow turbulence characteristics and the cascade geometry. Hanson's program, BBCASCADE, is incorporated into BFaNS, wherein it is applied to the rotor and stator in a stripwise manner. The spectra of upstream and downstream sound powers radiated by each strip are summed to obtain the total upstream and downstream sound-power spectra. The self-noise contributions are calculated by S. A. L. Glegg's theory, which is also applied in a stripwise manner. The current version of BFaNS is limited to fans with subsonic tip speeds.
Aerospace Engineering 2001-11-01AERONOV0111/1/2001
Safeguarding aircraft systems How several systems suppliers, airframers, and researchers have approached the need for increased system safeguards and protection against lightning strikes. One last look back A review of the technology news, discussions, and announcements from the first Aerospace Congress & Exhibition held September 10-13 in Seattle, WA. Powering unmanned aircraft Pratt & Whitney Canada is developing and testing a commercial turbofan engine derivative for high-altitude unmanned air vehicle applications. 737-800 winglet integration A joint venture between Boeing and Aviation Partners Inc. is involved in developing a performance improvement for in-service aircraft.
Optimal and Near-Optimal Take-off Maneuvers in the Presence of Windshear2000-01-559310/10/2000
This paper is concerned with the development of short-range forward-look guidance strategies for take-off maneuvers in the presence of windshear, with reference to flight in the vertical and horizontal plane. In addition to the horizontal shear, the presence of a downdraft is assumed. First, two open-loop optimal escape strategies are considered, viz. a performance maintaining strategy and a survival strategy, where it is assumed that the windshear is detected in an early stage of the encounter. Next, two closed-loop guidance strategies featuring short-range forward-look detection, which closely approximate the open-loop optimal solutions, are developed. Numerical results show that lateral maneuvering, in which an aircraft is turned away from the microburst, improves the flight performance significantly in both optimal escape strategies. The simulation results demonstrate that the employed closed-loop guidance strategies are able to closely approximate the open-loop optimal trajectories, in both longitudinal and lateral escape maneuvers.
Elferink, K., Visser, H.G.
Software Predicts Noise From Rotor-Wake/Stator InteractionsTBMG-696310/1/2000
V072 is a computer program that predicts the noise generated by the interaction between the rotor-blade wake (fan wake) and the stator vanes of a turbofan engine. The program is a combination of subroutines from prior NASA programs, with many added features.
Development of Aircraft-Specific Accident Investigation Reference Data Handbooks2000-01-21034/11/2000
This paper discusses the background, development, features and utility of aircraft-specific mishap investigation reference materials.
Gwynne, Donald G.
Passive Aircraft Anti Icing System Using Waste Heat9815427/13/1998
A passive anti icing system for the engine cowl of the Global Hawk UAV was designed and evaluated under a NASA Phase I SBIR Program. The system utilizes five Loop Heat Pipes to remove 3.8 kW of waste heat from the hydraulic system and deliver it to the engine inlet during icing conditions. During non-icing conditions the heat is bypassed to relief valves on the fuel ullage tank to maintain their temperature above freezing. The system masses 18.5 kg and replaces some 15 lbs of existing equipment. A Phase II program to integrate the system aboard the aircraft is underway with the goal of starting flight tests in early 1999.
Phillips, A. L. (Fred), Gernert, Nelson J.
Preliminary Sizing of a Supersonic Commercial Transport Between Mach 2.0 and 2.496558910/1/1996
To review specifications of a next-generation supersonic commercial transport (SCT), conceptual design is performed. The study covers the cruise speed between Mach 2.0 and 2.4, and the range between 4,000 nmi and 6,000 nmi with the number of seats fixed to 300. Engines for Mach 2.0, 2.2 and 2.4 are conceptually designed and the data are applied to Japan Aircraft Development Corporation (JADC)'s SCT sizing and performance prediction tool (“CAD”). The maximum take off weights (MTOW) are surveyed to find the best combination of cruise speed and range.
Takasu, Toru, Maekawa, Shoji, Ugai, Takashi, Mizuno, Hiroshi
Recommended Values of Rain and Hail Concentrations to be Considered in the Design of Turbine Engines96559610/1/1996
Extreme rain and hail conditions have caused engine rundowns during commercial airplane flights. In some cases, rain and hail conditions have caused all of the engines on the airplane to rundown simultaneously. This paper presents an analysis that quantifies the rain and hail threat in terms of the probability of encountering a storm of a given intensity for a given period of time. This analysis was used to establish recommended rain and hail concentration levels for design and certification of turbine engines.
Moravec, Bradford A., Patnoe, Michael W.
The Development and Application of Aircraft Radar Cross Section Prediction Methodology96553910/1/1996
In the last twenty years, radar cross section (RCS) reduction has emerged as an important and achievable way to improve the survivability of military aircraft. Specifically, the first manned aircraft to achieve a tactically significant reduction in RCS was the Lockheed Have Blue technology demonstrator, which was designed starting in 1975 and first flew in 1977. This was followed by a true military aircraft, the F-117 A, which became operational in 1983. Several very low observable (VLO) aircraft have been designed since then. All of them benefit from significant advances in the ability to predict and improve the RCS of an aircraft in its early design stages. The present paper describes these advances and their impact on aircraft design, using examples of reduced RCS aircraft designed before and after the techniques were evolved.
Aronstein, David C.
High Bypass Separate Flow Exhaust System Improved Thrust Efficiency By Modifying the Aft Centerbody9326229/1/1993
Scale model data indicated a performance dip in nozzle thrust efficiency at the cruise design conditions for a high bypass separate flow derivative exhaust system. A performance improvement program was implemented to determine the cause of the reduction in thrust efficiency and to identify viable redesigns. Data analysis indicated the performance loss to be a result of high Mach number fan and core exhaust interactions adversely affecting the pressure force on the exhaust system centerbody. Scale model testing confirmed the cause of performance loss and demonstrated that modifying the aft centerbody to reduce the plume Mach number recovers the loss in exhaust system thrust efficiency.
Babbitt, Bob
Prediction of Engine Casing Temperature of Fighter Aircraft for Infrared Signature Studies9209614/1/1992
Prediction of Engine Casing Temperature of Combat Aircraft for various operating conditions needs modeling of convective and radiative heat flow from the jet pipe to the casing. Radiation modeling which implies modeling of multiple reflections and scattering is done here by the radiance-irradiance approach. The engine layout is approximated by three co-axial cylinders and axially discretized. The shape factors (which account for the directional dependence of radiation heat flow) for various configurations needed in radiation modeling are analytically obtained by a sequence of shape factor algebra from standard configurations. The resulting non-linear algebraic equations of combined convection and radiation heat flow are solved by an iterative technique to give the engine casing temperatures which may be used for infrared signature studies.
Mahulikar, Shripad Prabhakar
Limit Cycle in the Longitudinal Motion of the USB STOL ASKA - Control System Functional Mockup and Actual Aircraft9210404/1/1992
The Japanese Quiet Short Take Off and Landing experimental aircraft named ASKA was developed and flight tested during 1977 till 1989. The control system hard and software were examined by the functional mock-up with using the actual hardware. The small longitudinal limit cycle was observed in the closed loop test when the Pitch Control Wheel Steering software was on in the mock-up testing. In this paper, first, the method to analyze and to expect the limit cycle based on the describing function was shown. The limit cycle was induced due to the nonlinearities in the automatic control mechanism. The nonlinearities in the hardware were examined to make the model to simulate the system on the computer. The method was shown effective to predict the limit cycle in the mock-up. Second, with using the flight measured dynamics, the limit cycle was concluded as on border line between existing and not, which coincides with the actual flight result. The functional mock-up and the describing function method can provide an accurate prediction of the limit cycle in the actual aircraft.
Uchikawa, Isoroku, Takesue, Youichi, Tsujimoto, Takashi, Uchida, Tadao, Okada, Noriaki, Ogawa, Toshio, Tada, Akira
Continuing the Etops Debate A Philosophy of Safety Overview9020219/1/1990
Five years ago, a new experiment in passenger air transportation for Atlantic Ocean crossings was introduced. It was given the acronym known as ETOPS and involves the development of passenger aircraft having only two-engines utilized for flights over such critical routes. The essence of the concern generated in this study is summarized by the safety analysis of Dr. Robert Besco, who states, “ETOPS is troublesome because the failure of only one system on an aircraft [In this case an engine] should not cause that aircraft to have to operate under emergency conditions.” Previously, a major cornerstone of civil aviation safety was the concept of providing multiple engine redundancy with a minimum of three-engines required for such critical flights. The accepted safety premise was that many unfavorable factors such as engine failures could jeopardize such flights and therefore the fail-safe method of providing three- or four-engine redundancy for these transports was prudent as well as comforting. Three- and four- engine transports are considered more safety efficient than two-engine transports or a possible one-engine transport of the future. Corporate, military and private flights in two- and one-engine aircraft fly such routes; however, they are not obligated to provide the highest safety standards that are required and expected of civil aviation transport operators. Formerly, two-engine transports were only permitted to fly over-land and over very short over-water segments primarily during domestic operations, and regulations still require that, when an engine fails, a landing must be made at the nearest suitable airport in point of time where a safe landing can be made. From over the United States and portions of Canada, this can often be accomplished in a relatively short time period of as little as thirty-minutes. Many emergency airfields are continuously available along such domestic routes. Indeed, in 1983 when one of these new twin-engine transports suffered a failure of both engines (due to fuel starvation problems), the aircraft glided powerless to a safe landing at a nearby emergency airfield. Such emergency airfields do not exist over the vast oceanic and desolate land areas over which ETOPS aircraft now may operate. The requirements for such emergency airfields were warved for ETOPS flights and instead, the - Federal Aviation Administration (FAA) gave credit for better engine reliability. Now, instead of requiring immediate landings of such two-engine transports after an engine failure, it is permissible to fly as long as three hours or cover over one-thousand miles, based heavily on the statistical success of continued favorable operation of the remaining engine. Proponents argue that safety is not compromised when engine redundancy is reduced in the ETOPS concept and certainly a good oceanic-track record during the next several decades is essential in order to promote the idea as sound. However, many aviation safety specialists are not convinced and cite several problem areas. For instance, some two-dozen engine shut-downs and over onehundred other serious situations have occurred where the ETOPS aircraft operated under emergency conditions or otherwise diverted to other destinations. In addition, some close calls have occurred, sirch as a one-engine diversion to an emergency airfield where the weather conditions had deteriorated to less than safe landing limits by the time the aircraft arrived. This was a very dangerous condition and one in which ETOPS aircraft are more vulnerable. During the summer of 1989 one trans-atlantic air-carrier had its ETOPS authority severely reduced due to the airlines' development of too many engine failures and related maintenance problems. Thus, aviation specialists continue to raise many and varied questions regarding ETOPS safety issues. It is unclear why these new generation engines with improved reliability were not initially fitfed onto new three- and four-engine air transport designs, thereby achieving the added safety benefits of both increased engine reliability as well as better engine redundancy. Other questions are detailed in the following text. This research paper contains historical as well as current information regarding the growing world trend in the use of two-engine civil air transports over vast desolate land areas and long oceanic routes.
Sprogis, Hal
Initial Service Experience With the Fokker 1008922389/1/1989
The paper outlines the behaviour of the Fokker 100 during its first year of commercial service with Swissair, the launch customer. Main design features are briefly discussed, together with selection criteria which prompted Swissair to become launch customer. Some aspects of crew training in relation to the introduction of glass cockpits are highlighted.
Ammann, H., van der Schraaf, A.
The Common/Same Type Rating: Human Factors and Other Issues8922299/1/1989
The concepts of the Common- and Same Type Rating are undoubtedly very attractive in today's cost-conscious airline operating environment. They allow the airlines to reduce their training and piloting costs while at the same time increasing scheduling flexibility by conducting mixed-fleet operations. However, certain types of flight deck differences beyond appropriate boundaries can result in increased workload and reduced crew performance. In addition, the introduction of advanced technologies, including those which enhance safety, can be delayed. This paper will attempt to shed some light on this complex issue by first defining what the Common/Same Type Rating is, by reviewing some of the experiences reported so far in mixed-fleet operations, and it will then discuss what some of the underlying human factors issues are and how known guidelines could be applied to future designs. The paper will conclude by trying to focus the industry's attention, (i.e., airplane manufacturers, airlines, and government agencies) on those areas in design, training, and scheduling where additional data is desperately needed.
Braune, R. J.
Application of KTRAN Transonic Small Disturbance Code to the Challenger Business Jet Configuration with Winglets88148310/1/1988
A Transonic Small Disturbance code originally developed at Canadair for the analysis of 3D wing/body/pylon/store configurations (AGARD CP-412-8) has been extended to calculate flows around complete aircraft such as the Challenger Executive Jet. The program uses a modified Transonic Small Disturbance equation discretized in cartesian and cylindrical coordinates and a grid embedding technique to capture flow details around specific components. The equation is solved using a successive line over-relaxation technique applied in two phases. In the first phase, the flow field is relaxed in the overall crude grid and in the winglet cylindrical grid. In the second phase, the crude grid and the various embedded fine grids are relaxed in alternating steps. The interaction between the various grids is through simple linear interpolation. The program is capable of representing Challenger-type wide body fuselages, large aspect ratio supercritical wings, high by-pass turbofan engines and canted winglets. Correlations with wind tunnel and flight test data demonstrate the capability of the method.
Kafyeke, F., Piperni, P., Robin, S.
Stability and Control Augmentation System of “ASKA”87233412/1/1987
“ASKA” is a STOL airplane with Upper Surface Blown type (USB) flaps used to perform research on powered lift technology developed by the National Aerospace Laboratory (NAL). ASKA has four high bypass ratio turbofan engines mounted above and forward of the wings, hydraulically actuated flight controls, and the Stability and Control Augmentation System (SCAS). The SCAS is a triple redundant system with three digital computers. In order to develop and evaluate its control laws, flight simulator tests have been conducted for 9 years during the design phase. Four flights have been devoted to evaluate functions of the SCAS and the control laws. The significant features of the control laws are to realize satisfactory flying qualities in the deep backside region at low airspeeds. As ASKA has not tested the flights in such regions up to present, this paper includes only the design features of the SCAS control laws, the results of the flight simulator tests, and interim outcomes of the flight tests on shallow USB flap configurations.
Kobayashi, Osamu, Tsujimoto, Takashi, Okada, Noriaki, Bando, Toshio
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