Showing papers on "Turbofan published in 1997"

Jet Propulsion: A Simple Guide to the Aerodynamic and Thermodynamic Design and Performance of Jet Engines

Abstract: Part I. Design of Engines for a New 600-Seat Aircraft: 1. The new large aircraft - requirements and background 2. The aerodynamics of the aircraft 3. The creation of thrust in a jet engine 4. The gas turbine cycle 5. The principle and layout of jet engines 6. Elementary fluid mechanics of compressible gases 7. Selection of bypass ratio 8. Dynamic scaling and dimensional analysis 9. Turbomachinery: compressors and turbines 10. Overview of the civil engine design Part II. Engine Component Characteristics and Engine Matching: 11. Component characteristics 12. Engine matching off-design Part III. The Design of the Engines for a New Fighter Aircraft: 13. A new fighter aircraft 14. Lift, drag and the effects of manoeuvring 15. Engines for combat aircraft 16. Design point for a combat aircraft 17. Combat engines off-design 18. Turbomachinery for combat aircraft Part IV. A Return to the Civil Engine: 19. A return to the civil transport engine 20. Conclusion.

Wave-Rotor-Enhanced Gas Turbine Engines

Abstract: The benefits of wave rotor topping in small (300- to 500-kW [400- to 700-hp] class) and intermediate (2000- to 3000-kW [3000- to 4000-hp] class) turboshaft engines, and large (350- to 450-kN [80,000- to 100,000-lb 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 topping is shown to enhance the specific fuel consumption and specific power of small- and intermediate-sized turboshaft engines significantly. The specific fuel consumption of the wave-rotor-enhanced large turbofan engine can be reduced while it operates at a 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 technical challenges.

Inlet Shape Effects on the Far-Field Sound of a Model Fan

Abstract: A wind tunnel test was conducted to determine the effects of inlet shape on fan radiated noise. Four inlet geometries, which included a long standard flight-type inlet, a short, aggressive flight inlet, a scarf inlet, and an elliptical inlet were investigated in the study. The fan model used in the study was a 0.1 scale of the Pratt and Whitney Advanced Ducted Propeller (ADP), an ultra high bypass ratio turbofan engine. Acoustic data are presented for a fan speed of 70 degrees (12,000) rpm) and a tunnel speed of 0.10 Mach number. The fan was configured with a 16-bladed rotor and a 40 stator vane set that were separated by 2.0 chord lengths. The radiated noise was measured with 15 microphones on a boom that traversed the length of the tunnel test section. Data from these microphones are presented in the form of sideline angle directivity plots. Noise associated with the test inlets was also predicted using a ray acoustics code. Inlet shape has been found to have a significant effect on both tone and broadband noise, and the non-axisymmetric inlet shape can be used for a noise reduction method.

Blockerless thrust reverser

Abstract: A blockerless thrust reverser for an aircraft having a podded nacelle housing a turbofan engine which produces core flow and a fan exit stream. Reverse thrust is obtained by diverting the fan exit stream into an annular slot formed in an outer wall of the nacelle where it is turned and discharged forwardly. The fan exit stream is directed into the annular slot by injecting high pressure streams of core flow into the fan exit stream at positions which are upstream of and adjacent to the annular slot. Reverse thrust is selectively obtained by a control means which selectively opens and closes the annular slot and the core jet injectors.

Development and Turbine Engine Performance of Three Advanced Rhenium Containing Superalloys for Single Crystal and Directionally Solidified Blades and Vanes

Abstract: Turbine inlet temperatures over the next few years will approach 1650°C (3000°F) at maximum power for the latest large commercial turbo fan engines, resulting in high fuel efficiency and thrust levels approaching 445 kN (100,000 lbs). High reliability and durability must be intrinsically designed into these turbine engines to meet operating economic targets and ETOPS certification requirements.This level of performance has been brought about by a combination of advances in air cooling for turbine blades and vanes, design technology for stresses and airflow, single crystal and directionally solidified casting process improvements and the development and use of rhenium (Re) containing high γ′ volume fraction nickel-base superalloys with advanced coatings, including full-airfoil ceramic thermal barrier coatings. Re additions to cast airfoil superalloys not only improve creep and thermo-mechanical fatigue strength but also environmental properties, including coating performance. Re dramatically slows down diffusion in these alloys at high operating temperatures.A team approach has been used to develop a family of two nickel-base single crystal alloys (CMSX-4® containing 3% Re and CMSX®−10 containing 6% Re) and a directionally solidified, columnar grain nickel-base alloy (CM 186 LC® containing 3% Re) for a variety of turbine engine applications. A range of critical properties of these alloys is reviewed in relation to turbine component engineering performance through engine certification testing and service experience.Industrial turbines are now commencing to use this aero developed turbine technology in both small and large frame units in addition to aero-derivative industrial engines. These applications are demanding, with high reliability required for turbine airfoils out to 25,000 hours, with perhaps greater than 50% of the time spent at maximum power. Combined cycle efficiencies of large frame industrial engines is scheduled to reach 60% in the U.S. ATS programme. Application experience to a total 1.3 million engine hours and 28,000 hours individual blade set service for CMSX-4 first stage turbine blades is reviewed for a small frame industrial engine.© 1997 ASME

Design of New Three Stage Low Pressure Turbine for the BMW Rolls-Royce BR715 Turbofan Engine

Abstract: In 1990, BMW and Rolls Royce plc (RR) joined to form a new company BWW-Rolls-Royce GmbH (BRR), to develop the BR700 family of engines aimed at the 12K and 25K lbs thrust range, using advanced technology and a modern organisation working in integrated teams to minimise the engine development timescales. After a successful development programme the BR710 engine rated at 14K lbs thrust, will shortly enter service in Gulfstream and Canadair Executive Jets.The recent launch of the BR715 engine at 21K lbs thrust, builds on the high pressure core developed for the BR710, plus a low pressure system with an increased diameter fan and 2 stage booster driven by a three stage turbine.This paper will describe, the advanced design technology incorporated, including the latest three dimensional aerodynamic philosophy using advanced high lift aerofoils for reduced parts count, plus the mechanical design issues addressed to optimise the LP turbine module configuration and the simultaneous design/make process employed to achieve the required parts delivery timescales.Copyright © 1997 by ASME

Active control of fan tones radiated from turbofan engines

Abstract: This paper continues an earlier companion paper in which the results of active control computer simulations aimed at assessing the potential for reducing the fan tones radiated from the inlet of turbofan engines were presented. The earlier study found that good levels of reductions in the sound power transmitted and sound pressure radiated towards the sidelines could be obtained by using error sensors located in the radiated far field. The present paper is also concerned with these control objectives, but with the important difference that the error sensors are located on the duct wall. The error sensor array described in this paper is comprised of a number of axial line array of discrete sensors uniformly distributed around the duct wall. By applying the appropriate phase delay between the sensor signals in a single axial array and summing, an estimate of the mode amplitudes versus axial propagation angle can be obtained. This in-duct pressure measurement can be related to the far-field radiation pattern...

The application of actuator disks to calculations of the flow in turbofan installations

Abstract: This paper discusses the application of an actuator disk model to the problem of calculating the asymmetric performance of a turbofan operating behind a nonaxisymmetric intake and due to the presence of the engine pylon. Good agreement between predictions and experimental results is demonstrated. Further validation of the model is obtained by comparison with the results of a three-dimensional calculation of an isolated fan operating with a nonaxisymmetric inlet. Some justification of the neglect of unsteady aspects of the flow in the fan is presented. The quantitative features of the interaction of the pylon and fan flow fields are discussed.

Reverse thrust inlet vortex inhibitor

Abstract: A method and apparatus are provided for inhibiting the formation of inlet vortices in a turbofan engine/nacelle installation 10 for use with an aircraft. The present invention redirects fan air from a fan air bypass duct 16, in a generally downward direction from the outer nacelle 18 in order to generate an air curtain 48 for inhibiting inlet vortex formation.

Development and evaluation of a coupled fan noise design system

Abstract: Fan wake/fan exit guide vane (FEGV) interaction tone noise has been known to be a major contributor to fan noise from typical turbofan engines. A Theoretical Fan Noise design/prediction System (TFaNS) has been developed which predicts tone noise from an inlet, fan, FEGV, and nozzle geometry. The system is excited by fan wakes impinging on the FEGV. The acoustic fields are coupled so as to permit waves to be reflected and transmitted throughout. A summary of the codes which comprise TFaNS is presented along with the coupling theory. Rotor/stator coupled tone noise predictions are made for a full scale engine and compared with data and with an earlier fan wake/FEGV interaction tone noise design system. The effects on the predictions of coupling to the-rotor are explored. A fully coupled fan noise prediction with an inlet, rotor, stator and nozzle is made for a fan rig and is compared with far-field directivity data. Results indicate that coupling is an important factor in turbofan noise prediction.

Interactive design tool for turbine based combined cycle engines

Abstract: A workstation-based interactive aerodynamic design tool has been developed to evaluate the performance and design constraints of a turbine based combined cycle engine concept for a hypersonic cruise vehicle. This tool uses one-dimensional aerodynamic and thermodynamic analysis techniques to model the inlet, turbine engine, ramjet burner, mixer ejector and single expansion ramp nozzle performance across the flight regime from sea level static to Mach 6.0 at altitude. The thermodynamic analysis may be performed using approximately thermally perfect or calorically perfect gasses. The design flight conditions, geometry, and parameters used in the analysis can be varied through a graphical user interface (GUI) and the change in engine performance is calculated and displayed immediately. A variety of graphical formats are used to present the results to the designer including numerical results, moving bar charts, and interactive schematic drawings. The tool provides printed output for performance plotting packages and has a restart capability. The GUI employs X based graphics widgets and the simulator runs on a single SGI workstation. The paper will detail the numerical methods used in the simulator and how it can been used in preliminary aerodynamic design. "Copyright c 1997 by the American Institute of Aeronautics and Astronautics, Inc. No copyright is asserted in the United States under Title 17, U.S. Code. The U.S. Government has a royalty-free license to exercise all rights under the copyright claimed herein for Governmental purposes. All other rights are reserved by the copyright owner. INTRODUCTION Recent advances in computer related technologies are changing the ways that engineers perform preliminary aerodynamic design studies. In the past, preliminary design was conducted using charts, tables and graphs of the performance of earlier similar configurations and final aerodynamic design was tested and verified using wind tunnel results. With the advent of large, powerful mainframe computers, some of the preliminary design tables and graphs could be numerically generated by solving the equations of motion and some of the final design results could be verified using computational fluid dynamics (CFD). Today's workstations and personal computers have computing power equal to that of the older mainframes. When coupled with a window operating systems and a graphical user interface, (GUI), a workstation can now be used to solve preliminary aerodynamic design problems interactively. This paper will describe the development of a software tool to perform preliminary aerodynamic design and evaluation of a propulsion system for a hypersonic cruise vehicle. The tool is based on previous flow solvers developed by the author, Refs. 1-4 , to perform preliminary aerodynamic design or educational analysis using interactive computer graphics on a single workstation. In Ref. 1 an interactive inlet design tool was developed to solve for the flow through external compression inlets. Through the use of a GUI, the designer could change the geometry, and the upstream and downstream flow conditions and immediately see the effects on inlet performance and drag. As the geometry and flow Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc. Figure 1: Turbine Based Combined Cycle (TBCC) propulsion system conditions were altered, the application would recompute the important flow variables and redisplay the geometry, shock wave locations, and output flow parameters. This package has been extended to also solve for the flow through rectangular mixed compression inlets, Ref. 2. In Ref. 3 some of the coding from the the inlet design program was used to produce an educational package to study simple turbojet engines. In this package, the user sets input design conditions using the GUI and the code performs a Brayton cycle analysis as outlined in Ref. 5 to determine thrust and specific fuel consumption. This package has also recently been extended to analyze turbofan and afterburning turbojet engines, Ref. 4. The propulsion system to be modeled in this application is a turbine based combined cycle (TBCC) engine for a hypersonic cruise vehicle. The conceptual vehicle would cruise at Mach 10 using a hydrogen burning scramjet propulsion system. Scramjet powered vehicles, however, must be accelerated by some other means until sufficient ram compression exists for full supersonic combustion. While a variety of propulsion systems have been proposed to accelerate the vehicle, this study employs conventional fueled turbojets which are integrated with the ramjet/scramjet flow path. The turbojets would be used for take-off, low speed cruise, ferry and, in conjunction with the ramjet, would accelerate the vehicle to near Mach 3. At this point sufficient ram compression exists to operate the scramjet flow path in a subsonic combustion "single throat" ramjet mode, Ref. 6. As Mach number then increases to approximately Mach 6, the engine transitions to full supersonic combustion. A major concern in using a combined tubojet/ramjet/scramjet propulsion system is the integration of the turbojet with the ramjet/scramjet flow path since the turbojet represents a weight and volume penalty at high speed cruise. Details of the conceptual design, shown in Fig. 1, and its predicted performance can be found in Ref. 7. In Fig. 1, a row of turbojets are placed in a bay in the fuselage above the ramjet flow path. The turbojets and ramjets share a common inlet and nozzle while an ejector is used to control the airflow split between them. The ejector flap can be used to close off the ramjet flowpath for low speed operation as shown in Fig. 2a. A thermal choke is achieved at the ramjet spraybar which acts as a low speed afterburner for the turbojet. Soon after take-off the ejector flap is raised as shown in Fig. 2b. The ejector allows more flow to pass through the ramjet throat as the airflow through the turbojet decreases with Mach and altitude. Near Mach 3, a two position inlet flap is lowered while the ejector flap is raised to completely isolate the tubojet from the high temperature ramjet/scramjet stream as shown in Fig. 2c. The ramjet then accelerates the system to Mach 6 for scramjet transition.

In-flight balancing of fan on turbofan jet engine

Abstract: An apparatus and a method for in-flight balancing of fan 10 on a turbofan jet engine, after a loss of blade 12, involves detecting an imbalance due to the loss of a blade, by means of a vibration monitor 50, optical sensors 56, a blade locator magnet (68 fig. 2), and coil (55), and a strain gauge (69 fig. 2). A computer 51 is linked via a radio link 54,62,60 to a computer 63, and a power switch 64 and voltage regulator 65 are provided. When a blade 12 is lost, the power switch 64 is caused to send electrical impulses to selected membrane valves 35, the membranes of which burn, melt or are ruptured by an explosive charge, whereby fluid communication between a tyre-shaped ring 30 and selective blade cavities 15, is effected to admit fluid to the selected cavities 15 in fan blades 12, adjacent to the place of the lost blade, to compensate for the lost mass. Fluid is supplied to the ring via a delivery line 22, centrifugal force then urging fluid into lines 34. Electrical power is supplied generated by coils (66 fig. 2) on the hub 11 rotating in relation to stationary magnets (67).

Advanced Low-Noise Research Fan Stage Design

Abstract: This report describes the design of the Advanced Low-Noise Research Fan stage. The fan is a variable pitch design, which is designed at the cruise pitch condition. Relative to the cruise setting, the blade is closed at takeoff and opened for reverse thrust operation. The fan stage is a split flow design with fan exit guide vanes (FEGVs) and core stators. The fan stage design is combined with a nacelle and engine core duct to form a powered fan/nacelle subscale model. This model is intended for use in combined aerodynamic, acoustic, and structural testing in a wind tunnel. The fan has an outer diameter of 22 in. and a hub-to-tip of 0.426 in., which allows the use of existing NASA fan and cowl force balance and rig drive systems. The design parameters were selected to permit valid acoustic and aerodynamic comparisons with the Pratt & Whitney (P&W) 17- and 22-in. rigs previously tested under NASA contract. The fan stage design is described in detail. The results of the design axisymmetric and Navier-Stokes aerodynamic analysis are presented at the critical design conditions. The structural analysis of the fan rotor and attachment is included. The blade and attachment are predicted to have adequate low-cycle fatigue life and an acceptable operating range without resonant stress or flutter. The stage was acoustically designed with airfoil counts in the FEGV and core stator to minimize noise. A fan/FEGV tone analysis developed separately under NASA contract was used to determine the optimum airfoil counts. The fan stage was matched to the existing nacelle, designed under the previous P&W low-noise contract, to form a fan/nacelle model for wind tunnel testing. It is an axisymmetric nacelle for convenience in testing and analysis. Previous testing confirmed that the nacelle performed as required at various aircraft operating conditions.

Predicted performance characteristics of a variable cycle turbofan

Abstract: The selective bleed turbofan is a two shaft, three compressor, variable cycle aircraft engine. Its bypass ratio (BPR) can be modulated to suit the operating flight conditions. At subsonic flight speeds it operates as a medium bypass turbofan (low pressure mode or LPM). It becomes a low bypass turbofan (high pressure mode or HPM) when flying faster and is capable of supersonic cruise without reheat (RH). The aim of this paper is to describe the selection of the main engine design parameters in the HP mode and to outline the challenges to be met to enable the powerplant to operate in the other mode. The components have to perform satisfactorily over a wide range of varying conditions. The resulting demands are reflected in many ways, for example in the operating lines of the low pressure compressors and in the overall efficiency changes with increased bypass ratio. In the case of the present investigation the two bypass ratios selected are 0.3 for the high pressure mode and 0.7 for the low pressure mode. The latter was selected after an examination of a choice of bypass ratios ranging from 0.4 to 1.0. The change in variable stators required to achieve the LP mode cycle is indicated. The effect of changing bypass ratio on engine performance and handling is shown. The main conclusion is that the design of this engine seems to be feasible within current technological capabilities and further investigation is encouraged because it appears to yield significant benefits.

Blade containment system for a turbofan engine

Abstract: Casing assembly for a gas turbine engine comprising an annular cross section casing around which a plurality of flexible material are wound. One or more rigid panels are positioned between the wound flexible material and the annular casing. The rigid panel provides enhanced containment of a detached fan blade or part of a fan blade by distributing the load of the detached blade across the wound flexible material.

A finite element formulation for aft fan duct acoustic radiation

Abstract: A finite element code has been developed for the prediction of the radiated acoustic field in and around the aft fan duct of a turbofan engine. The acoustic field is modeled based on the assumption that the steady flow in and around the nacelle is irrotational as is the acoustic perturbation. The geometry of the nacelle is axisymmetric and the acoustic source is harmonic and decomposed into its angular harmonics. The steady flow is computed on the acoustic mesh and provides data for the acoustic calculations. The jet is included in the steady flow potential flow model by separating the interior and exterior flow outside the aft fan duct with a thin barrier created by disconnecting the computational domain. The jet and exterior flow is allowed to mix at a defined distance downstream. In the acoustic radiation model continuity of acoustic particle velocity is implicitly satisfied across the shear layer by careful treatment of the surface integral which appears in the FEM formulation. Pressure continuity is enforced by using a penalty constraint on the shear layer. A reliable frontal solution routine which originally involved extensive I/O operations to minimize core storage has been updated to eliminate most of the inefficient direct access reading and writing with considerable impact on computational time and is now found to be competitive with banded LU solvers. Example calculations are given which show the success achieved in satisfying the complicated interface conditions on the shear layer and the characteristics of the solutions at relatively high frequencies where the refinement of the mesh is a limiting consideration for practical computations.

Analysis of Windmilling Characteristics for a Twin-Spool Turbofan Engine

Abstract: The windmilling characteristics of twin-spool, high bypass ratio turbofan engine have been analyzed. This analysis is an extension of the previously reported analysis for a single-spool turbojet engine. As before, the aerodynamic performances of engine components are determined by incorporating the available cascade loss correlations. For a given flight condition, the steady-state windmilling conditions are determined by iteratively balancing the mass flow rate and angular momentum through the two spools. Compared to the turbojet analysis, the new analysis requires determination of bypass ratio and work split between the two spools. Some of the calculation results have been compared against the limited data available for a CF-6 engine, and the two show good agreement. The present method is thus shown to be capable in predicting turbofan engine’s windmilling characteristics during its design stage.Copyright © 1997 by ASME

Two stage mixer ejector for turbofan noise suppression

Abstract: A two-stage mixer ejector concept ("TSMEC") (10) is disclosed for suppressing the noise emanted by jet aircraft. This TSMEC was designed to help older engines (12) meet the stringent new federal noise regulations in the United States of America, known as "Stage III". In an illustrated embodiment, the TSMEC comprises: a lobed engine nozzle (18) attached to the rear of a turbofan engine (12); a short shroud (20) that straddles the exit end of the engine nozzle; two rings (30, 32) of convergent/divergent primary and secondary lobes within the engine nozzle shroud; and a ring (26) of accurate gaps that preceed the shroud and the second nozzle ring. The lobes (30,32) are complimentary shaped to rapidly mix the turbine's hot exhaust flow with cooler air including entrained ambient air, at supersonic speed. This drastically reduces the jet exhaust's velocity and increases jet mixing, thereby reducing the jet noise.

Turbo-fan jet engine

Abstract: The fan jet engine system has an ejector nozzle at the trailing end of the engine at a bypass outlet. An annular hollow zone is in front of the ejector nozzle to form a pneumatic connection with the air suction openings at the fan engine or other parts of the aircraft.The pneumatic cavities of the ejector nozzle and the front suction part are in regulated connection, for the purpose of total optimisation of the by-pass flow

A variable cycle jet engine for ASTOVL aircraft

Abstract: Describes a novel concept in aircraft propulsion: investigates a variable cycle jet engine for a supersonic advanced short take‐off vertical landing (ASTOVL) aircraft. The engine is the selective bleed turbofan. The selective bleed turbofan is a two shaft, three compressor, variable cycle gas turbine. 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. A preliminary design of an ASTOVL aircraft from Cranfield, the S‐95, was used as the vehicle. Outlines the performance of the engine and its integration with the aircraft. Explains off‐design engine performance characteristics and describes variable geometry requirements. The major advantage of this engine is that all the components are employed all the time, for all operating modes, thus incurring low weight penalties. Predicts that the aircraft/ engine combination will perform in a satisfactory way, meeting most performance targets provided that some improvements are carried out.

Active noise control of low speed fan rotor-stator modes

Abstract: This report describes the Active Noise Cancellation System designed by General Electric and tested in the NASA Lewis Research Center's 48 inch Active Noise Control Fan. The goal of this study was to assess the feasibility of using wall mounted secondary acoustic sources and sensors within the duct of a high bypass turbofan aircraft engine for active noise cancellation of fan tones. The control system is based on a modal control approach. A known acoustic mode propagating in the fan duct is cancelled using an array of flush-mounted compact sound sources. Controller inputs are signals from a shaft encoder and a microphone array which senses the residual acoustic mode in the duct. The canceling modal signal is generated by a modal controller. The key results are that the (6,0) mode was completely eliminated at 920 Hz and substantially reduced elsewhere. The total tone power was reduced 9.4 dB. Farfield 2BPF SPL reductions of 13 dB were obtained. The (4,0) and (4,1) modes were reduced simultaneously yielding a 15 dB modal PWL decrease. Global attenuation of PWL was obtained using an actuator and sensor system totally contained within the duct.

Interactive Educational Tool for Turbofan and Afterburning Turbojet Engines

Abstract: A workstation-based, interactive educational computer program has been developed at the NASA Lewis Research Center to aid in the teaching and understanding of turbine engine design and analysis. This tool has recently been extended to model the performance of two-spool turbofans and afterburning turbojets. The program solves for the flow conditions through the engine by using classical one-dimensional thermodynamic analysis found in various propulsion textbooks. Either an approximately thermally perfect or calorically perfect gas can be used in the thermodynamic analysis. Students can vary the design conditions through a graphical user interface; engine performance is calculated immediately. A variety of graphical formats are used to present results, including numerical results, moving bar charts, and student-generated temperature versus entropy (Ts), pressure versus specific volume (pv), and engine performance plots. The package includes user-controlled printed output, restart capability, online help screens, and a browser that displays teacher-prepared lessons in turbomachinery. The program runs on a variety of workstations or a personal computer using the UNIX operating system and X-based graphics. It is being tested at several universities in the midwestern United States; the source and executables are available free from the author.

'cfd for the aerodynamic design of bombardier's global express® high performance jet"

Abstract: In this paper the studies that led to the development of the aerodynamic configuration of the Global Express are described. Emphasis is put on the methodology, in particular the role of CFD and wind tunnel testing in the design of the transonic wing and its high-lift system, and in the integration of the turbofan engines with the aft fuselage.

Innovative concept for cascade thrust reverser without blocker doors

Abstract: This paper reports the progress being made on an innovative thrust reverser concept for potential use on commercial turbofan aircraft. The concept employs a unique application of fluidics by using the readily available high-pressure engine core flow to divert the much larger fan flow out through a cascade opening. In our simulated 5:1 bypass ratio fan engine, we have been able to turn 100% of the fan flow into the cascade opening by injecting a mass flow that is less than 3% of the fan flow. This amount of core flow bleed is well within current engine capabilities. With a simple cascade vane design, we have achieved more than 40% thrust reversal. This performance is comparable to conventional blocker door configurations without suffering the corresponding weight, mechanical complexity, and aerodynamic cruise penalties. This report describes our experimental and analytic study, and states our conclusions on the effects of various geometric and aerodynamic parameters.

An experimental study of unsteady forcing in the f-109 turbofan engine

Abstract: Phase-locked, unsteady, surface pressure measurements were made across the swept stator vanes of the single stage of axial compression in a running AlliedSignal F-109 turbofan engine. Measurements were nominally made at seven vane-chord pressure-side and suction-side locations for two engine-axis radial positions at 90, 94, and 97 percent engine speed. The acquired pressure responses were ensemble averaged, time resolved, and decomposed into steady data and unsteady amplitude, frequency, and phase data. Analysis of the decomposed, unsteady, fundamental frequency data produced position-vs.phase maps which demonstrate ambiguous unsteady-disturbance propagation direction. In addition, significant variation between the data taken at the separate engine-axis positions along the swept vanes is observed. Arguments are presented to support the assertion that the phase propagation direction ambiguity combined

Mach wave elimination in supersonic jets

Abstract: Experimental results are presented on a method that eliminates Mach waves from the exhaust of supersonic jets and, hence, that removes a strong component of supersonic jet noise. Elimination is achieved by surrounding the jet with an annular stream at prescribed velocity and temperature so that all turbulent motions become intrinsically subsonic. No mechanical suppressors are used. Implementation of the technique in a typical turbofan engine is estimated to increase takeoff thrust with minimal impact on overall fuel consumption.