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Showing papers on "Turbofan published in 1977"


Patent
11 Mar 1977
TL;DR: In this paper, the cooling capacity of an existing engine fuel/oil heat exchanger is supplemented by an existing internal combustion engine (i.e., turbofan) to reduce the size of the fan air/oil exchanger and minimize an aircraft engine performance penalty.
Abstract: This invention serves to cool the lubricant used by a mechanical constant speed drive (CSD) driving an aircraft alternator which aircraft is powered by a turbofan engine. The system cooling capacity is supplemented by an existing engine fuel/oil heat exchanger thereby effectuating a reduction in size of the CSD fan air/oil heat exchanger and minimizing an aircraft engine performance penalty.

85 citations


01 Jan 1977
TL;DR: Weight and dimensional relationships that are used in aircraft preliminary design studies are analyzed and all estimating relations stem from physical principles, not statistical correlations.
Abstract: Weight and dimensional relationships that are used in aircraft preliminary design studies are analyzed. These relationships are relatively simple to prove useful to the preliminary designer, but they are sufficiently detailed to provide meaningful design tradeoffs. All weight and dimensional relationships are developed from data bases of existing and conceptual turbofan engines. The total propulsion system is considered including both engine and nacelle, and all estimating relations stem from physical principles, not statistical correlations.

67 citations


Patent
Peter Kent1
23 May 1977
TL;DR: In this paper, a cylindrical cowl is used to cover the engine core with an annular shaped cavity and ports in the cowl allow air to be brought into the cavity from the bypass ducting of the engine.
Abstract: This invention relates to apparatus which harmlessly purges all fluids which leak from the fittings of a turbofan aircraft engine. Purging is achieved by passing a low speed flow of ventilating air along the exterior walls of the engine combustor section. This is accomplished by enshrouding the engine core with a cylindrical cowl. A space between the cowl and the engine combustor makes an annular shaped cavity. Ports in the cowl allow air to be brought into the cavity from the bypass ducting of the engine. The rapidly moving volume of air thus brought in vaporizes any fuel which leaks into the cavity. The vaporized air/fuel mixture is then exhausted into the hot gas plume emitted from the rear of the core engine.

55 citations


Patent
01 Feb 1977
TL;DR: In this paper, a closed loop control that closes the loop on the pressure ratio across the fan by adjusting fuel flow and/or exhaust nozzle area in a turbofan, variable exhaust nozzle installation is proposed.
Abstract: Increased thrust during the transonic flight mode of an aircraft powered by a gas turbine engine is realized by a closed loop control that closes the loop on the pressure ratio across the fan by adjusting fuel flow and/or exhaust nozzle area in a turbofan, variable exhaust nozzle installation. Fan pressure ratio is scheduled as a function of corrected fan speed and actual pressure ratio provides an error signal to readjust engine operation to null out this error signal. A turbine inlet temperature limit signal is generated to prevent inadvertent overheating and it or this pressure ratio error signal is selected for providing the lower fuel flow value.

43 citations


Patent
02 Aug 1977
TL;DR: In this article, a propulsion system for use primarily in V/STOL aircraft is provided with a variable cycle, double bypass gas turbofan engine and a remote augmenter to produce auxiliary lift.
Abstract: A propulsion system for use primarily in V/STOL aircraft is provided with a variable cycle, double bypass gas turbofan engine and a remote augmenter to produce auxiliary lift. The fan is oversized in air-pumping capability with respect to the cruise flight requirements of the remainder of the engine and a variable area, low pressure turbine is capable of supplying varying amounts of rotational energy to the oversized fan, thereby modulating its speed and pumping capability. During powered lift flight, the variable cycle engine is operated in the single bypass mode with the oversized fan at its maximum pumping capability. In this mode, substantially all of the bypass flow is routed as an auxiliary airstream to the remote augmenter where it is mixed with fuel, burned and exhausted through a vectorable nozzle to produce thrust for lifting. Additional lift is generated by the high energy products of combustion of the variable cycle engine which are further energized in an afterburner and exhausted through a thrust vectorable nozzle at the rear of the engine. In the cruise operating mode, the fan is driven at a slower rotational speed and all of the bypass flow is directed to a variable area bypass injector where it is mixed with the variable cycle engine products of combustion and exhausted as a mixture through the rear nozzle. The variable cycle engine can be selectively operated in the single or double bypass mode during cruise to maximize efficiency as the aircraft speed is varied. Because the total installed thrust is available in the forward cruise mode, an aircraft powered by the propulsion system of the present invention has substantially greater capabilities in the areas of excess power, acceleration time and combat ceiling when compared to prior art conceptions embodying separate lift plus lift/cruise engines. Additionally, by not having to develop and maintain a separate lift engine system, a substantially lower aircraft lift cycle cost is possible.

40 citations


Journal ArticleDOI
TL;DR: In this article, an improved prediction system for direct combustion noise is discussed, where the effects of fuel nozzle number and burner length are used to predict the acoustic power level, peak frequency, and full-scale engine acoustic transmission loss due to combustion noise.
Abstract: The development of an improved prediction system for direct combustion noise is discussed. Expressions for acoustic power level, peak frequency, and full-scale engine acoustic transmission loss due to combustor/duct coupling and turbine attenuation are derived in terms of readily available performance and geometry parameters from the burner and turbine. New parameters introduced by the prediction system include the effects of fuel nozzle number and burner length. Predictions are in good agreement with noise data obtained from component rig tests on several JT8D burner configurations, and full-scale turbofan engines. The applicability of the system to the prediction of combustion noise levels, spectra, and directivity from full-scale engines is demonstrated for four PW i.e., the JT8D-109, JT9D-7A, JT9D-70, and the prototype JT10D.

30 citations


Patent
28 Jun 1977
TL;DR: In this paper, a thrust reverser nozzle for coaxial-flow turbofan engines comprising target-type deflector doors which are hinged for deployment about a fixed axis by means of actuation about single fixed pivots mounted on support structure on either side of the engine nacelle rearward portion.
Abstract: A thrust reverser nozzle for coaxial-flow turbofan engines comprising target-type deflector doors which are hinged for deployment about a fixed axis by means of actuation about single fixed pivots mounted on support structure on either side of the engine nacelle rearward portion. The deflector door's outer surfaces are shaped to match existing aerodynamic contours of the aircraft engine nacelle so as to provide a lower boattail angle for improved drag characteristics in the stowed or normal flight position. In that position, the deflector door interior configuration comprises a portion upstream of the engine exhaust nozzle exit plane and a downstream "fishmouth" portion through which flows hot engine exhaust gases surrounded circumferentially by cool air discharged from the engine fan. Geometry of the stowed fishmouth is sized and shaped to take advantage of mixing and shearing action between the exhaust streams so as to produce a desired variable area nozzle effect on the engine operation and thereby improve its forward thrust performance. Geometry of the upstream portion of the inside surface of the doors is sized and shaped with end plates so that when the deflector doors move to the deployed position, the exhaust streams are diverted outward and forward to produce a desired level of reverse thrust.

29 citations


01 Feb 1977
TL;DR: In this article, the effects of exhaust nozzle geometry on radiation patterns of low frequency noise at the source were evaluated using scale model tests, and the results were used to improve the component prediction techniques derived under the FAA-RD-74-125, III.
Abstract: : This program was directed towards elements of combustor and turbine noise; the latter including turbine tone interaction with jet stream turbulence Combustor (Core) Noise - Investigations were conducted to determine the variables affecting source strength, spectrum shape, and farfield directivity This investigation include scale model tests to evaluate the effects of exhaust nozzle geometry on radiation patterns of low frequency noise A full-scale combustor rig test was used to identify the controlling variables of combustor noise at the source Two engine tests were run to validate the findings from the scale model tests and add to the overall data base of core noise measurements The relationship between combustor source noise and emissions was studied and qualitative trends identified for advanced low emissions combustors Turbine Noise - Studies were made of the attenuation of high frequency turbine noise by downstream blade rows, the broadband noise generation by turbines, and the controlling parameters for turbine tone/jet stream interaction This included a turbine rig test in single and multistage configurations, along with a unique data acquisition system Scale model tests were used to define the effect of the pertinent aero-acoustic parameters on turbine tone scattering by jet stream turbulence The results of these investigations were used to improve the component prediction techniques derived under the Core Engine Noise Control Program (FAA-RD-74-125, III) These improved prediction techniques were used to predict the noise contribution of each source for high bypass turbofan engines representative of current and advanced technology (Author)

23 citations


Journal ArticleDOI
TL;DR: In this paper, a static jet noise experiment with five primary flow nozzles were used with a common secondary nozzle to simulate exhaust flows of turbofan engines with bypass ratios from 1 to 5.
Abstract: Mixing of primary and secondary flows in a conventional turbofan engine provides a means of reducing jet noise. By shaping the nozzle exit velocity profile, noise reduction greater than that resulting from fully mixed flow has been achieved. In a static jet noise experiment, five primary flow nozzles were used with a common secondary nozzle to simulate exhaust flows of turbofan engines with bypass ratios from 1 to 5. Data are shown which relate jet noise to the location, extent, and magnitude of the peak velocity region. In general, minimum noise is obtained for inverted profiles where the outer area peak velocity is 5. to 15% greater than the reference uniformly mixed velocity, and the area of the peak velocity region is 40 to 50% of the total flow area. The inverted flow profiles produce noise characteristics similar to multielement jet suppressor nozzles, i.e., low frequencies are reduced and high frequencies are increased. It is shown that these spectral effects can be used to obtain a balanced noise signature.

23 citations


Journal ArticleDOI
TL;DR: In this article, the core noise from a YF-102 high bypass ratio turbofan engine was investigated through the use of simultaneous measurements of internal fluctuating pressures and far field noise.
Abstract: Core noise from a YF-102 high bypass ratio turbofan engine was investigated through the use of simultaneous measurements of internal fluctuating pressures and far field noise. Acoustic waveguide probes, located in the engine at the compressor exit, in the combustor, at the turbine exit, and in the core nozzle, were employed to measure internal fluctuating pressures. Spectra showed that the internal signals were free of tones, except at high frequency where machinery noise was present. Data obtained over a wide range of engine conditions suggest that below 60% of maximum fan speed the low frequency core noise contributes significantly to the far field noise.

23 citations


01 Jan 1977
TL;DR: In this paper, the aerodynamic design methodology for these models is discussed and some of the preliminary test results are presented which indicate that propeller net efficiencies near 80 percent were obtained for high disk loading propellers operating at Mach 0.8.
Abstract: The increased emphasis on fuel conservation in the world has stimulated a series of studies of both conventional and unconventional propulsion systems for commercial aircraft. Preliminary results from these studies indicate that a fuel saving of 14 to 40 percent may be realized by the use of an advanced high-speed turboprop. This turboprop must be capable of high efficiency at Mach 0.8 cruise above 9.144 km altitude if it is to compete with turbofan powered commercial aircraft. Several advanced aerodynamic concepts were investigated in recent wind tunnel tests under NASA sponsorship on two propeller models. These concepts included aerodynamically integrated propeller/nacelles, area ruling, blade sweep, reduced blade thickness and power (disk) loadings several times higher than conventional designs. The aerodynamic design methodology for these models is discussed. In addition, some of the preliminary test results are presented which indicate that propeller net efficiencies near 80 percent were obtained for high disk loading propellers operating at Mach 0.8.

01 Feb 1977
TL;DR: In this paper, an improved method for predicting both direct and indirect combustion noise from aircraft engines is developed and experimentally evaluated by conducting rig experiments and by comparing with data from several full scale engines.
Abstract: : Improved methods for predicting both direct and indirect combustion noise from aircraft engines are developed and experimentally evaluated by conducting rig experiments and by comparing with data from several full scale engines. Comparison of predictions with full scale engine data indicated that direct combustion noise is the dominant source for the P and WA engines investigated. The direct combustion noise prediction system includes expressions for acoustic power level, peak frequency and full-scale engine acoustic transmission loss due to combustor/duct coupling and turbine attenuation. These expressions are derived in terms of readily available performance and geometry parameters from the burner and turbine. New parameters introduced by the prediction system include the effects of fuel nozzle number and burner length. Predictions are shown to be in good agreement with data obtained from component rig tests on several JT8D type burner configurations (including single and multiple fuel nozzle, conventional and low emission designs). In addition, when transmission losses are accounted for, the predictions are also shown to be in good agreement with observed combustion noise levels and peak frequencies from four P and WA turbofan engines (i.e. the JT8D-109, JT9D-7A, JT9D-70 and the prototype JT10D). Predicted combustion noise directivity patterns and spectra shapes are determined empirically, using the data from both the rig tests and these four engines. Results from the analytical and experimental combustion noise investigations are used to identify combustion noise reduction methods obtainable through modifications in burner design and/or performance parameters.


Patent
John F. Hurley1
01 Sep 1977
TL;DR: In this paper, the temperature sensing element of the fuel control unit is bathed in a continuously fresh sample of ambient air drawn in from the engine inlet by an air-jet pump that is powered by pressurized air bled-off from the bypass fan duct.
Abstract: This invention relates to apparatus which provides a positive flow of ambient air through the temperature sensor of the fuel control equipment used with a turbofan engine. The temperature sensing element of the fuel control unit is bathed in a continuously fresh sample of ambient air drawn in from the engine inlet by an air-jet pump that is powered by pressurized air bled-off from the bypass fan duct.

Patent
27 Jan 1977
TL;DR: In this article, a modified conventional lobe mixer is utilized to invert flow streams in a turbofan engine to reduce the noise generated by the gaseous streams in the coaxial streams of the engine.
Abstract: Jet noise generation occasioned from the gaseous streams in the coaxial streams of a turbofan engine when they as well as the ambient stream encounter is reduced by designing the engine so that the value of the true velocity of the outer stream is substantially higher than the value of the true velocity of the inner stream. It is contemplated that for a turbofan engine the fan stream and primary or engine core streams are inverted so that the higher velocity stream would be in the outer coaxial passage immediately upstream of the point where the streams discharge to ambient. A modified conventional lobe mixer is utilized to invert flow streams.

Patent
Joseph Raphael Kasmarik1
31 Jan 1977
TL;DR: A multiplicity of retractable spoilers are placed at spaced angular intervals within the annular duct which carries the bypass airstream in a turbofan engine as discussed by the authors, which results in a decrease in thrust from the engine without any change in throttle setting.
Abstract: A multiplicity of retractable spoilers are placed at spaced angular intervals within the annular duct which carries the bypass airstream in a turbofan engine. Deployment of the spoilers into the bypass annular duct causes a partial blockage of the airstream. This results in a decrease in thrust from the engine without any change in throttle setting.

Patent
27 Dec 1977
TL;DR: In this paper, an approach for internally mixing fan exhaust with primary exhaust forward of the nozzle exit plane of a turbofan engine having an outer engine cowling, a splitter wall structure and a tail plug disposed in mid engine upstream of the exit plane is provided.
Abstract: Apparatus for internally mixing fan exhaust with primary exhaust forward of the nozzle exit plane of a turbofan engine having an outer engine cowling, a splitter wall structure and a tail plug disposed in mid engine upstream of the nozzle exit plane is provided by a plurality of struts disposed radially between the engine tail plug and the outer engine cowling upstream of the nozzle exit plane and a plurality of vortex generating means disposed on the struts for generating vorticies across the entire engine exhaust area.

Patent
26 Apr 1977
TL;DR: In this article, a thrust augmented twin spool turbofan engine system is used to control instability at high altitude, low Mach number conditions wherein an engine electronic control monitors engine inlet temperature, engine burner pressure and Mach number.
Abstract: A thrust augmented twin spool turbofan engine system to control instability at high altitude, low Mach number conditions wherein an engine electronic control monitors engine inlet temperature, engine burner pressure and Mach number, and inhibits fuel flow to the outermost augmentor segment when the engine inlet temp drops below 25° F, when engine burner pressure drops below 120 psia or when the Mach number drops below 0.4. When fuel flow to the outermost augmentor is inhibited, upmatch logic modifies the variable nozzle AJ trim signal closed 8% and the power level angle trim signal up 4°.

Patent
25 Oct 1977
TL;DR: In this article, a short length afterburner assembly for a jet propulsion engine having a fan bypass includes cold and hot air cross-over passages and a plurality of flame stabilization swirler vanes associated with a balanced load controller for positioning the vanes parallel to hot gas stream flow from a jet engine core.
Abstract: A short length afterburner assembly for a jet propulsion engine having a fan bypass includes cold and hot air cross-over passages and a plurality of flame stabilization swirler vanes associated with a balanced load controller for positioning the vanes parallel to hot gas stream flow from a jet engine core when the afterburner is off and in an inclined position to such gas stream flow when fuel is injected therein during afterburner operation thereby to produce flame spread within the afterburner core by a combination of translatory and swirling motions; and wherein atomized fuel for afterburner combustion is injected into hot gases ducted through hot air cross-over passages from the jet engine core to produce premix and prevaporization of fuel upstream of fixed flameholders and wherein cold fan bypass air cross-over passages have movable turbulator grids positioned during afterburner operation for mixing cold air flow with the bypassed hot core gas at the fixed flameholders during afterburner operation and wherein the turbulator grids are positioned parallel to gas flow when the afterburner is not in operation by means of the balanced load controller to produce a balanced variable geometry mechanical load to that on the flame stabilization swirler vanes.

Patent
09 Dec 1977
TL;DR: In this article, the normal compressor stator schedule of a fan speed controlled turbofan engine is temporarily varied to substantially close the stators to increase the fuel flow and compressor speed in order to maintain fan speed and thrust.
Abstract: Upon a landing approach, the normal compressor stator schedule of a fan speed controlled turbofan engine is temporarily varied to substantially close the stators to thereby increase the fuel flow and compressor speed in order to maintain fan speed and thrust. This running of the compressor at an off-design speed substantially reduces the time required to subsequently advance the engine speed to the takeoff thrust level by advancing the throttle and opening the compressor stators.

Book
Fred Teren1
01 Jan 1977
TL;DR: In this article, the minimum time accelerations of aircraft turbofan engines are calculated using a piecewise linear engine model and an algorithm based on nonlinear programming, which allows such trajectories to be readily calculated on a digital computer with a minimal expenditure of computer time.
Abstract: Minimum time accelerations of aircraft turbofan engines are presented. The calculation of these accelerations was made by using a piecewise linear engine model, and an algorithm based on nonlinear programming. Use of this model and algorithm allows such trajectories to be readily calculated on a digital computer with a minimal expenditure of computer time.

01 Jun 1977
TL;DR: In this paper, an extended Kalman filter is used to provide the best estimate of the state of the turbofan engine based on currently available sensor outputs, which allows continuing control of the engines in the event of a sensor failure.
Abstract: In this paper, a failure detection and correction strategy for turbofan engines is discussed. This strategy allows continuing control of the engines in the event of a sensor failure. An extended Kalman filter is used to provide the best estimate of the state of the engine based on currently available sensor outputs. Should a sensor failure occur the control is based on the best estimate rather than the sensor output. The extended Kalman filter consists of essentially two parts, a nonlinear model of the engine and up-date logic which causes the model to track the actual engine. Details on the model and up-date logic are presented. To allow implementation, approximations are made to the feedback gain matrix which result in a single feedback matrix which is suitable for use over the entire flight envelope. The effect of these approximations on stability and response is discussed. Results from a detailed nonlinear simulation indicate that good control can be maintained even under multiple failures.

01 Sep 1977
TL;DR: A real time, hybrid computer simulation of a turbofan engine is described and shown to match the steady state and transient performance of the engine over a wide range of flight conditions and power settings.
Abstract: A real time, hybrid computer simulation of a turbofan engine is described. Controls research programs involving that engine are supported by the simulation. The real time simulation is shown to match the steady state and transient performance of the engine over a wide range of flight conditions and power settings. The simulation equations, FORTRAN listing, and analog patching diagrams are included.

Patent
25 Nov 1977
TL;DR: A thrust reverser for a turbo-fan engine is provided with a sleeve (28) which normally forms part of the fan outlet duct When reverse thrust is required the sleeve is pushed rearwards by means of a jack and shuts off the duct outlet.
Abstract: A thrust reverser for a turbo-fan engine is provided with a sleeve (28) which normally forms part of the fan outlet duct When reverse thrust is required the sleeve is pushed rearwards by means of a jack and shuts off the duct outlet An annulus (37) in the sleeve (28) houses a number of rings which form a cascade of blades (34) These blades are exposed to the air flow by the rearward movement of the sleeve (28) The air from the fan is then deflected by the cascade of blades and consequently exerts a reverse thrust on the aircraft

01 Oct 1977
TL;DR: A series of iterative combustor pressure rig tests were conducted on two combustor concepts applied to the AiResearch TFE731-2 turbofan engine combustion system for the purpose of optimizing combustor performance and operating characteristics consistant with low emissions as mentioned in this paper.
Abstract: A series of iterative combustor pressure rig tests were conducted on two combustor concepts applied to the AiResearch TFE731-2 turbofan engine combustion system for the purpose of optimizing combustor performance and operating characteristics consistant with low emissions. The two concepts were an axial air-assisted airblast fuel injection configuration with variable-geometry air swirlers and a staged premix/prevaporization configuration. The iterative rig testing and modification sequence on both concepts was intended to provide operational compatibility with the engine and determine one concept for further evaluation in a TFE731-2 engine.

Journal ArticleDOI
Fred Teren1
TL;DR: In this article, minimum-time accelerations of the F100 turbofan engine are presented, where a piecewise linear engine model, having three state variables and four control variables, is used to obtain the minimumtime solutions.
Abstract: Minimum-time accelerations of the F100 turbofan engine are presented. A piecewise-linear engine model, having three state variables and four control variables, is used to obtain the minimum-time solutions. The linear model which applies at a given time in the trajectory is determined by calculating a normalized distance from the current state to the equilibrium state associated with each linear model. The linear model associated with the closest equilibrium point is then used. The control histories for the minimum-time solutions are used as input to a nonlinear simulation of the F100 engine to verify the accuracy of the piecewise linear solutions.

01 Dec 1977
TL;DR: In this article, a baseline compressor test stage was designed as well as a candidate rotor and two candidate stators that have the potential of reducing endwall losses relative to the baseline stage.
Abstract: A baseline compressor test stage was designed as well as a candidate rotor and two candidate stators that have the potential of reducing endwall losses relative to the baseline stage. These test stages are typical of those required in the rear stages of advanced, highly-loaded core compressors. The baseline Stage A is a low-speed model of Stage 7 of the 10 stage AMAC compressor. Candidate Rotor B uses a type of meanline in the tip region that unloads the leading edge and loads the trailing edge relative to the baseline Rotor A design. Candidate Stator B embodies twist gradients in the endwall region. Candidate Stator C embodies airfoil sections near the endwalls that have reduced trailing edge loading relative to Stator A. Tests will be conducted using four identical stages of blading so that the designs described will operate in a true multistage environment.

01 Aug 1977
TL;DR: In this article, the dynamic stability and control characteristics of a four-engine turbofan transport model having an upper-surface blown jet flap were investigated by means of the free-flight technique in the Langley full-scale tunnel.
Abstract: The dynamic stability and control characteristics of a four-engine turbofan transport model having an upper-surface blown jet flap were investigated by means of the free-flight technique in the Langley full-scale tunnel. The flight characteristics of the model were investigated through a range of lift coefficients from 3 to 8 with all four engines operating and with one outboard engine not operating. Static characteristics were investigated by conventional power-on force tests over the flight-test angle-of-attack range and through the stall.

Proceedings ArticleDOI
27 Mar 1977
TL;DR: In this article, the authors describe the NASA Lewis Research Center's engine fan noise facility, where fan models are electrically driven to 20,600 RPM in either the inlet mode or exhaust mode to facilitate study of both fore and aft fan noise.
Abstract: Acoustical and mechanical design features of NASA Lewis Research Center's engine fan noise facility are described. Acoustic evaluation of the chamber, which is lined with an array of stepped wedges, is described. Results from the evaluation in terms of cut-off frequency and non-anechoic areas near the walls are detailed. Fan models are electrically driven to 20,600 RPM in either the inlet mode or exhaust mode to facilitate study of both fore and aft fan noise. Inlet noise characteristics of the first fan tested are discussed and compared to full-scale levels. Turbulence properties of the inlet flow and acoustic results are compared with and without a turbulence reducing screen over the fan inlet.

Journal ArticleDOI
TL;DR: In this paper, aeroacoustic tests of model coannular nozzles have shown that less noise is generated if the higher velocity jet is exhausted from the outer annular passage rather than from the primary nozzle.
Abstract: Recent aeroacoustic tests of model coannular nozzles have shown that less noise is generated if the higher velocity jet is exhausted from the outer annular passage rather than from the primary nozzle. These findings are of particular significance to the duct burning turbofan engine being studied for application to an advanced supersonic transport airplane. Unlike conventional turbofan engines that have peak velocities from the primary nozzle, it is possible to design a DBTF engine to have a fan velocity higher than that of the primary flow. In this paper are presented the results of a model test program that covers a range of fan to primary area ratios from 0.75 to 1.2, and a range of fan-to-primar y velocity ratios from 0.4 to 2.8. Correlations are presented that relate radiated sound power to fan velocity, fan-to-primary velocity ratio, and fan-to-primary area ratio. Corresponding exhaust plume velocity traverse data are presented which suggest that the observed noise benefits may be because of the more rapid decay of the annular flow caused by shear stresses on the inner surface which result from the lower velocity primary flow.