Showing papers on "Turbofan published in 1995"

Active control of aircraft engine inlet noise using compact sound sources and distributed error sensors

Abstract: An active noise control system using a compact sound source is effective to reduce aircraft engine duct noise. The fan noise from a turbofan engine is controlled using an adaptive filtered-x algorithm. Single, multi channel control systems are used to control the fan blade passage frequency (BPF) tone and the BPF tone and the first harmonic of the BPF tone for a plane wave excitation. The multi channel control system is used to control fan tones and a high pressure compressor BPF tone simultaneously, and any spinning mode. A compact sound source is employed to generate the control field. This compact sound source consists of an array of identical thin, cylindrically curved panels (125) with an inner radius of curvature corresponding to that of the engine inlet. These panels are flush mounted inside the inlet duct (Inlet Wall) and sealed on all edges to prevent leakage around the panel.

Improved Finite Element Modeling of the Turbofan Engine Inlet Radiation Problem

Abstract: Improvements have been made in the finite element model of the acoustic radiated field from a turbofan engine inlet in the presence of a mean flow. The problem of acoustic radiation from a turbofan engine inlet is difficult to model numerically because of the large domain and high frequencies involved. A numerical model with conventional finite elements in the near field and wave envelope elements in the far field has been constructed. By employing an irrotational mean flow assumption, both the mean flow and the acoustic perturbation problem have been posed in an axisymmetric formulation in terms of the velocity potential; thereby minimizing computer storage and time requirements. The finite element mesh has been altered in search of an improved solution. The mean flow problem has been reformulated with new boundary conditions to make it theoretically rigorous. The sound source at the fan face has been modeled as a combination of positive and negative propagating duct eigenfunctions. Therefore, a finite element duct eigenvalue problem has been solved on the fan face and the resulting modal matrix has been used to implement a source boundary condition on the fan face in the acoustic radiation problem. In the post processing of the solution, the acoustic pressure has been evaluated at Gauss points inside the elements and the nodal pressure values have been interpolated from them. This has significantly improved the results. The effect of the geometric position of the transition circle between conventional finite elements and wave envelope elements has been studied and it has been found that the transition can be made nearer to the inlet than previously assumed.

Inlet acoustic mode measurements using a continuously rotating rake

Abstract: Comprehensive measurements of the spinning acoustic mode structure in the inlet of the advanced ducted propeller were obtained using a unique method that was first proposed by Sofrin. A continuously rotating microphone system was employed. Three inlet configurations with cut-on as well as cut-off stator vane sets were tested. The cut-off stator was designed to suppress all modes at the blade passing frequency. Rotating rake measurements indicate that several extraneous circumferential modes, possibly due to the interaction between the rotor and small interruptions in the casing tip treatment, were present. The cut-on stator produced the expected circumferential modes plus higher levels of the unexpected modes seen with the cut-off stator. HE next generation of fan engines will likely employ a marriage of turbofan and propeller technologies to achieve significant noise and fuel consumption reductions. The ad- vanced ducted propeller (ADP) model used in this investi- gation was designed and built by Pratt and Whitney, a Di- vision of United Technologies, and tested in the NASA Lewis 9- by 15-ft Anechoic Wind Tunnel. Typical of propeller tech- nology, the ADP allows for the in-flight adjustment of the blade pitch angle. This provides reverse thrust and optimum performance over a wide range of conditions. The duct pro- vides the noise suppression advantage of a conventional fan engine. Since future engines are expected to use still higher bypass ratios, fan noise is likely to be the dominant engine source. One of the most important features of fan tone noise is its modal structure. Knowledge of these spinning modes helps to identify the generation mechanism, control duct propa- gation (thus, mode knowledge is needed for acoustic treat- ment design) and control far-field radiation. Previous at- tempts at direct mode measurements1'3 have faced formidable practical difficulties such as: very large axial and circumfer- ential arrays of wall microphones that are not practical for the short ducts of ultrahigh bypass engines, and radial mea- surements upstream of the fan that introduce a wake that interacts with the rotor, thus causing extraneous modes. A continuously rotating microphone technique first proposed by Sofrin4 overcomes the problem of wake-generate d modes, reduces the number of microphones and the duct length re- quired. This technique has been implemented for the first time in this investigation. Two important features of this tech- nique are as follows:

Inlet acoustic mode measurements using a continuously rotating rake

Abstract: Comprehensive measurements of the spinning acoustic mode structure in the inlet of the advanced ducted propeller were obtained using a unique method that was first proposed by Sofrin. A continuously rotating microphone system was employed. Three inlet configurations with cut-on as well as cut-off stator vane sets were tested. The cut-off stator was designed to suppress all modes at the blade passing frequency. Rotating rake measurements indicate that several extraneous circumferential modes, possibly due to the interaction between the rotor and small interruptions in the casing tip treatment, were present. The cut-on stator produced the expected circumferential modes plus higher levels of the unexpected modes seen with the cut-off stator. HE next generation of fan engines will likely employ a marriage of turbofan and propeller technologies to achieve significant noise and fuel consumption reductions. The ad- vanced ducted propeller (ADP) model used in this investi- gation was designed and built by Pratt and Whitney, a Di- vision of United Technologies, and tested in the NASA Lewis 9- by 15-ft Anechoic Wind Tunnel. Typical of propeller tech- nology, the ADP allows for the in-flight adjustment of the blade pitch angle. This provides reverse thrust and optimum performance over a wide range of conditions. The duct pro- vides the noise suppression advantage of a conventional fan engine. Since future engines are expected to use still higher bypass ratios, fan noise is likely to be the dominant engine source. One of the most important features of fan tone noise is its modal structure. Knowledge of these spinning modes helps to identify the generation mechanism, control duct propa- gation (thus, mode knowledge is needed for acoustic treat- ment design) and control far-field radiation. Previous at- tempts at direct mode measurements1'3 have faced formidable practical difficulties such as: very large axial and circumfer- ential arrays of wall microphones that are not practical for the short ducts of ultrahigh bypass engines, and radial mea- surements upstream of the fan that introduce a wake that interacts with the rotor, thus causing extraneous modes. A continuously rotating microphone technique first proposed by Sofrin4 overcomes the problem of wake-generate d modes, reduces the number of microphones and the duct length re- quired. This technique has been implemented for the first time in this investigation. Two important features of this tech- nique are as follows:

Multi-spool by-pass turbofan engine

Abstract: A multi-spool turbofan engine (6) has a high pressure spool (26) that operates at high idle RPM so as to power accessories and minimize acceleration time from idle to full rated RPM without producing high idle thrust which compromises the landing maneuver and ground operation. Gas flow from a combustor (40) is split immediately aft of a high pressure turbine (36) so that only a portion thereof passes through intermediate and low pressure turbine blades (22 and 24) on a low pressure spool (12), thereby reducing significantly the power developed by the low pressure spool (12). Accordingly, the RPM of the fan (18) is significantly reduced thereby reducing the thrust of the engine (6) at idle conditions. Moreover, the relatively high idle RPM of the high pressure spool (26) is maintained due to bleed off of combustion gas immediately aft of the high pressure turbine (36). The relatively high idle RPM for the high pressure spool (26) results in adequate power being delivered to an alternator (56).

Calculation of Two-Phase Flow in Gas Turbine Combustors

Abstract: A method is presented for computing steady two-phase turbulent combusting flow in a gas turbine combustor. The gas phase equations are solved in an Eulerian frame of reference. The two-phase calculations are performed by using a liquid droplet spray combustion a model and treating the motion of the evaporating fuel droplets in a Lagrangian frame of reference. The numerical algorithm employs nonorthogonal curvilinear coordinates, a multigrid iterative solution procedure, the standard k-{epsilon} turbulence model, and a combustion model comprising an assumed shape probability density function and the conserved scalar formulation. The trajectory computation of the fuel provides the source terms for all the gas phase equations. This two-phase model was applied to a real piece of combustion hardware in the form of a modern GE/SNECMA single annular CFM56 turbofan engine combustor. For the purposes of comparison, calculations were also performed by treating the fuel as a single gaseous phase. The effect on the solution of two extreme situations of the fuel as a gas and initially as a liquid was examined. The distribution of the velocity field and the conserved scalar within the combustor, as well as the distribution of the temperature field in the reaction zone and in themore » exhaust, were all predicted with the combustor operating both at high-power and low-power (ground idle) conditions. The calculated exit gas temperature was compared with test rig measurements. Under both low and high-power conditions, the temperature appeared to show an improved agreement with the measured data when the calculations were performed with the spray model as compared to a single-phase calculation.« less

Wave rotor-enhanced gas turbine engines

Abstract: The benefits of wave rotor-topping in small (400 to 600 hp-class) and intermediate (3000 to 4000 hp-class) turboshaft engines, and large (80,000 to 100,000 lb(sub f)-class) high bypass ratio turbofan engines are evaluated. Wave rotor performance levels are calculated using a one-dimensional design/analysis code. Baseline and wave rotor-enhanced engine performance levels are obtained from a cycle deck in which the wave rotor is represented as a burner with pressure gain. Wave rotor-toppings is shown to significantly enhance the specific fuel consumption and specific power of small and intermediate size turboshaft engines. The specific fuel consumption of the wave rotor-enhanced large turbofan engine can be reduced while operating at significantly reduced turbine inlet temperature. The wave rotor-enhanced engine is shown to behave off-design like a conventional engine. Discussion concerning the impact of the wave rotor/gas turbine engine integration identifies tenable technical challenges.

UHB Engine Fan Broadband Noise Reduction Study

Abstract: A study has been completed to quantify the contribution of fan broadband noise to advanced high bypass turbofan engine system noise levels The result suggests that reducing fan broadband noise can produce 3 to 4 EPNdB in engine system noise reduction, once the fan tones are eliminated Further, in conjunction with the elimination of fan tones and an increase in bypass ratio, a potential reduction of 7 to 10 EPNdB in system noise can be achieved In addition, an initial assessment of engine broadband noise source mechanisms has been made, concluding that the dominant source of fan broadband noise is the interaction of incident inlet boundary layer turbulence with the fan rotor This source has two contributors, ie, unsteady life dipole response and steady loading quadrupole response The quadrupole contribution was found to be the most important component, suggesting that broadband noise reduction can be achieved by the reduction of steady loading field-turbulence field quadrupole interaction Finally, for a controlled experimental quantification and verification, the study recommends that further broadband noise tests be done on a simulated engine rig, such as the GE Aircraft Engine Universal Propulsion Simulator, rather than testing on an engine statically in an outdoor arena The rig should be capable of generating forward and aft propagating fan noise, and it needs to be tested in a large freejet or a wind tunnel

Advanced Turbine Systems annual program review

Abstract: Integrated High Performance Turbine Engine Technology (IHPTET) is a joint Air Force, Navy, Army, NASA, ARPA, and industry program focused on developing turbine engine technologies, with the goal of doubling propulsion capability by around the turn-of-the-century, and thus providing smaller, lighter, more durable, more affordable turbine engines in the future. IHPTET`s technology development plan for increasing propulsion capability with respect to time is divided into three phases. This phased approach reduces the technological risk of taking one giant leap, and also reduces the {open_quotes}political{close_quotes} risk of not delivering a product for an extended period of time, in that the phasing allows continuous transfer of IHPTET technologies to our warfighters and continuous transfer to the commercial sector (dual-use). The IHPTET program addresses the three major classes of engines: turbofan/turbojet, turboshaft/turboprop, and expendables.

Advanced turbofan blade refurbishment technique

Abstract: The purpose of the work reported here is to investigate whether the lessons learned from the work of Suder et al. can be used to reduce the in-service performance deterioration of a fan on a high bypass ratio turbofan engine. To this end, a back-to-back test was done on the fan of an RB211-22B engine with the cooperation of Delta Airlines. The fan and engine were first overhauled per normal airline practice and cell-tested to establish that the engine performance met flight acceptance standards. This test, which the engine passed, also established a performance baseline for the overhauled engine. At this point the fan blade leading edge had not been filed or scraped and the blade surfaces had not been polished because the leading edge damage and blade surface roughness fell within the acceptable limits specified by the manufacturer for normal overhaul practice. After the cell test, the fan was removed from the engine and sent to Sermatech International where the following additional operations were performed: (1) the blade surfaces were polished to a finish of 20 rms {micro}in; (2) leading edge roughness due to particle impact damage was removed and the leading edge was polished to a finish ofmore » 20 rms {micro}in; (3) the leading edge shape was rounded and the leading edge thickness was reduced over the first 5--10% of chord. Test results indicated a 0.7% drop in thrust specific fuel consumption (lb fuel/lb thrust/hr) relative to the baseline engine after the enhanced fan overhaul. Based on the results of Suder et al. (1995) it appears that 70--80% of this performance gain is due to the thin smooth leading edge and the remainder to the highly polished finish of the blade.« less

Axisymmetric Aerodynamic Numerical Analysis of a Turbofan Engine

Abstract: An equation and numerical solution procedure for a steady-state aerodynamic analysis of a turbofan engine in the meridional plane is presented. An axisymmetric average of the three-dimensional fluid flow equations within the engine is presented. The equations simultaneously represent the major engine components and include terms for blade forces, loss, combustor heat addition, blockage, bleeds and convective mixing in the mixer. These general equations are specialized with modeled terms and advanced to a steady-state with a time-marching scheme on a multi-block grid. The procedure is verified with comparisons to one-dimensional analytic solutions and a compressor stage. How this procedure can capture component interactions is demonstrated with a solution for General Electric’s Energy Efficient engine, which is representative of modern turbofan engines.Copyright © 1995 by ASME

Experiments on the Active Control of a Turbofan Inlet Noise Using Compact, Lightweight Inlet Control and Error Transducers

Abstract: In this paper the active noise control of turbofan tonal inlet noise using lightweight, compact control sound sources and distributed pressure sensors mounted in the inlet wall is demonstrated on an operational Pratt and Whitney JT15D engine. Furthermore, by the use of an optimal selection of the control configuration, control spillover problems observed in previous experiments were eliminated and global reductions of up to 5 dB of the blade passage frequency achieved with an electrical power consumption of approximately 10 watts. The work thus demonstrates that active control of turbofan inlet noise using inlet control transducers is certainly feasible for aerospace applications and achievable in practice with realistic transducers.

Method of disassembling a gas turbine engine power plant

Abstract: A method and apparatus which facilitate separation of a gas turbofan powerplant into modules for shipping, maintenance and repair is disclosed. Various construction details are developed which provide means for mounting a fan cowling to an engine core in a manner which permits transference of operational loads from the fan cowling to the engine core and separation of the fan cowling from the engine core. In one embodiment, a fan cowling (46) is attached to an engine core (18) by a plurality of radially extending through struts (64). The through struts include a bolted joint (72) which permits separation of a powerplant (12) into a first module and a second module. For this embodiment, a method for varying between an assembled and disassembled condition is comprised of manipulating the joints between an engaged and disengaged position and axially moving the separate modules along a longitudinal centerline (14).

Fan Noise Reduction From a Supersonic Inlet During Simulated Aircraft Approach

Abstract: A series of experiments was conducted to investigate the radiation of fan noise from a supersonic inlet during a simulated aircraft approach. A scaled-down model of an axisymmetric, mixed-compression, supersonic inlet (P-inlet) was used in conjunction with a 10.4 cm (4.1 in) diameter turbofan engine simulator as the noise source. The tests were conducted at an outdoor facility under static conditions. The main goal of the experiment was to reduce the forward radiating fan noise by modifying the auxiliary inlet doors. The modified doors are designed to reduce the inlet distortion to the fan face. In addition, the new door design also uses a converging flow passage in order to take advantage of the noise attenuation due to the choking effect at the auwiliary door. The simulator was tested at 60 percent design speed in an attempt to match the simulator noise source to that of a real aircraft engine on approach. Both aerodynamic and acoustic measurements were taken in the experiments. The results show that when compared to the original design, the modified auxiliary inlet doors reduced the circumferential inlet distortion to the fan face by a factor of two. The key result is that the blade passing frequency tone has been decreased by an average of 6 dB in the forward sector for the modified door design. Results from the closed auxiliary inlet door case are also presented to provide additional comparisons

Transition Control and Performance of the Selective Bleed Variable Cycle Turbofan

Abstract: The selective bleed turbofan is a two shaft, three compressor, variable cycle aircraft engine. At subsonic flight speeds it operates as a medium bypass turbofan. It becomes a low bypass turbofan when flying faster and is capable of supersonic cruise in the dry mode.The aim of this paper is to outline the requirements and response of the engine during the transition phase from one mode of operation to the other. Thus it is follow-on work of Ref. 7. To achieve this several control devices can be employed. These are the fuel flow, the variable compressor stators, the nozzle areas and the valve areas of the bypass ducts.The preliminary analysis described in this paper indicates that all the criteria of performance are satisfied and that the transition can be carried out successfully provided the control variables are co-ordinated properly.Copyright © 1995 by ASME

A New Approach to the Design of the Large Turbofan Power Plant

Abstract: The lower direct operating costs of the Big Twin subsonic transports encourage the building of ever larger turbofan engines installed on the wings. The steadily improving reliability of the turbofa...

An Advanced Control System for Turbofan Engine: Multivariable Control and Fuzzy Logic (Application to the M88-2 Engine)

Abstract: The search for better performance of present and future turbofan engine involves an increase on the number of variable geometries and thus of control loops.As we can not or do not want to disregard the interaction between loops any more, the future control systems will therefore be multivariable. The aim of the architecture of multivariable control presented here is to optimize a performance index during transients.This architecture consists of an inner loop which optimizes the performance index taking in account the limitations, an outer loop which brings the nominal steady-state offsets to zero and a trajectory which allows to take into account the topping schedule limitation.This basic architecture can be improved by fuzzy supervisor.Indeed, two control outputs are generated according to the description above:- the first one optimizes the thrust and does not care very much about LP stall margin limitation,- the second one optimizes again the thrust and strongly takes low pressure stall margin limitation into account.The fuzzy logic then allows to do a compromise between these two control outputs according to the engine state.Simulation results showing the efficiency of the method are given.Copyright © 1995 by ASME

Multivariable Control of the F-100 Turbofan Engine Using Edmunds' Method

Abstract: This paper presents the design of a F-100 turbofan engine controller using the multivariable frequency domain method, called the Edmunds' method. The idea of this method is to design a simple-structure controller so that a desired performance can be achieved. The parameters of the controller are adjusted so that the closed-loop frequency response is as closed as possible, in a least square sense, to a desired response. This paper emphasizes the choice of the desired target performance which greatly affects the successfiilness of the Edmunds' method. In this paper, an operating point (sea level, static, PLA = 52°) is selected as a design point, and a controller is designed at this point. First, a linear state-space model obtained from a nonlinear engine model is transformed into a transfer function. Based on the information of this transfer function, the desired target closed-loop transfer function and structure of the controller can be determined. Next, the Edmunds' method is used to design a controller. Finally, the performance of the controller is evaluated by computer simulation and performances of different controllers are compared.

Simulating Indirect Thrust Measurement Methods for High-Bypass Turbofans

Abstract: As yet, there is no known reliable method for directly measuring the thrust of a turbofan in flight. Manufacturers of civil turbofans use various indirect thrust measurements to indicate performance of an engine to the flight deck. Included among these are: Engine Pressure Ratio (EPR), Integrated Engine Pressure Ratio (IEPR), Fan mechanical speed (N 1 ), and various Turbine Gas Temperatures such as ITT or EGT. Of key concern is whether these thrust indicators give an accurate account of the actual engine thrust. The accuracy of these methods, which is crucial at take-off may be compromised by various types of common engine deterioration, to the point where a thrust indicator may give a false indication of the health and thrust of the engine. A study was done to determine the effect of advanced engine cycles on typical values of these parameters. A preliminary investigation of the effects of common kinds of turbofan deterioration was conducted to see how these faults can affect both actual engine performance and the indirect thrust indicators

Electric motor air-cooling device for turbo fan unit

Abstract: PURPOSE: To provide a motor air-cooling device for a turbofan unit, for performing air-cooling of an electric motor to avoid damage to the motor. CONSTITUTION: This air-cooling fan 32 which creates a secondary air stream by means of a common electric motor 10 is provided and only the secondary air stream is urged to flow into the electric motor 10 through inlets 28, 30 separated from a main air stream created by the impeller 20 of a turbofan unit driven at speeds above 50,000 rpm. After the secondary air stream flows therethrough, it is mixed with the main stream and released to the atmosphere through an outlet 35.

Computation of noise radiation from turbofans: A parametric study

Abstract: This report presents the results of a parametric study of the turbofan far-field noise radiation using a finite element technique. Several turbofan noise radiation characteristics of both the inlet and the aft ducts have been examined through the finite element solutions. The predicted far-field principal lobe angle variations with duct Mach number and cut-off ratio compare very well with the available analytical results. The solutions also show that the far-field lobe angle is only a function of cut-off ratio, and nearly independent of the mode number. These results indicate that the finite element codes are well suited for the prediction of noise radiation characteristics of a turbofan. The effects of variations in the aft duct geometry are examined. The ability of the codes to handle ducts with acoustic treatments is also demonstrated.

Surge detector using engine signature

Abstract: PURPOSE: To facilitate detection of a surge condition by obtaining derivatives of engine operation characteristics such as fan rotation speed and temperature of exhaust gas, comparing the derivatives with threshold values, respectively, and increasing and decreasing a count of a counting means based on the results to judge a surge condition based on the count. CONSTITUTION: A turbo fan engine 1 consisting of a housing 2, a fan 3, a compressor 4, a combustor 5, and a turbine 6 is provided with a surge detection device 10 in which engine operation characteristics such as a fan speed signal by means of a fan speed converter 7 and an exhaust gas temperature signal by means of a temperature converter 8 are input. Derivatives of each engine operation characteristic are obtained here. A count is increased only when derivatives of a first engine operation characteristic exceed a first threshold value and/or derivatives of a second engine operation characteristic exceed a second threshold value, and a count is decreased in other cases. When the count is a predetermined value larger than 1, it is judged that a surge condition occurs, and a fuel flow rate into the combustor 5 is controlled.

Thrust deflecting device for turbofan engine

Abstract: PURPOSE:To provide a thrust deflecting device for a turbofan engine which can deflect thrust with good responsiveness and without changing the area of an exhaust nozzle. CONSTITUTION:Deflecting nozzle movable parts 1-4 are provided in the circumferential direction of the bypass passage of the exhaust nozzle 5 of a turbofan engine so as to position them at right angles with each other. The movable parts 1-4 can be displaced in the radial direction of the exhaust nozzle 5, and the movable parts opposite to each other on a diagonal line are displaced in the opposite directions so as to deflect thrust. Consequently the thrust can be deflected without actually changing the total area of the exhaust nozzle.

A new approach to the design of the large turbofan power plant

Abstract: The three major large turbofan engines-GE CF6, PW JT9 and RR RB211 have been in service flying the jumbo jets now for well over 20 years. Their creation made an enormous advance in transport aircraft propulsion efficiency and reduction of noise. They transformed world travel. This was not achieved without major problems and mechanical failures in service during the early years. The combination of large physical size, high pressure ratio cycles, high turbine entry temperatures, and engine/aircraft installation incompatibilities were collectively the cause of many of these problems. The aircraft constructors and airline operators were saying that the engine builders had been too optimistic and that never again would they countenance the launch of a new aircraft design powered by a new unproven engine design. The kind of serious problems and mechanical failures experienced were : turbine and fan disc failures related to new materials; fan blade detachment ; engine casing distortion; separation of the complete engine from the wing mounting ;

Design Considerations of a Versatile Simulator for High-Bypass Turbofans

Abstract: The continuing requirement for performance improvement of higher thrust turbofans is met by increased bypass ratios. The trend toward higher bypass ratios and relatively large-diameter low-pressure-ratio fans requires innovative design approaches, which include shorter inlets, slimmer nacelles, shorter fan ducts and exhaust systems, and possible elimination of thrust reverser. The success of this new generation of high-bypass ducted turbofans depends on understanding the acoustic impact from reduced treatment areas, inlet-fan coupling and operability, as well as overall system performance. To achieve these goals a versatile scale model propulsion simulator large enough to operate as a fan rig, yet small enough to be installed in a wind runnel for evaluating overall acoustic, operability, and system performance, was developed. The criteria of designing such a simulator and its capabilities are discussed in this paper