US20050247046A1 - Jet nozzle mixer - Google Patents

Jet nozzle mixer Download PDF

Info

Publication number
US20050247046A1
US20050247046A1 US11/027,982 US2798205A US2005247046A1 US 20050247046 A1 US20050247046 A1 US 20050247046A1 US 2798205 A US2798205 A US 2798205A US 2005247046 A1 US2005247046 A1 US 2005247046A1
Authority
US
United States
Prior art keywords
stage
lobes
mixer
stage mixer
tubular housing
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Abandoned
Application number
US11/027,982
Inventor
Jack Anderson
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Individual
Original Assignee
Individual
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Individual filed Critical Individual
Priority to US11/027,982 priority Critical patent/US20050247046A1/en
Publication of US20050247046A1 publication Critical patent/US20050247046A1/en
Abandoned legal-status Critical Current

Links

Images

Classifications

    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02KJET-PROPULSION PLANTS
    • F02K1/00Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
    • F02K1/46Nozzles having means for adding air to the jet or for augmenting the mixing region between the jet and the ambient air, e.g. for silencing
    • F02K1/48Corrugated nozzles
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02KJET-PROPULSION PLANTS
    • F02K1/00Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
    • F02K1/46Nozzles having means for adding air to the jet or for augmenting the mixing region between the jet and the ambient air, e.g. for silencing
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02KJET-PROPULSION PLANTS
    • F02K1/00Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
    • F02K1/54Nozzles having means for reversing jet thrust
    • F02K1/64Reversing fan flow
    • F02K1/70Reversing fan flow using thrust reverser flaps or doors mounted on the fan housing
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02KJET-PROPULSION PLANTS
    • F02K3/00Plants including a gas turbine driving a compressor or a ducted fan
    • F02K3/02Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
    • YGENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y10TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10TTECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T29/00Metal working
    • Y10T29/49Method of mechanical manufacture
    • Y10T29/49346Rocket or jet device making

Definitions

  • the present invention relates to jet nozzle mixers for aircraft jet engines and, in particular, to improvements in effecting a greater cooling and a lower noise level in exhaust gases emanating from such engines and in increasing power and fuel efficiency.
  • Noise (decibel) level in jet aircraft engines is established by laws and regulations, specifically promulgated by the International Civil Aviation Organization (ICAO), Annex 16. At present, commercial jet aircraft weighing over 75,000 pounds (34,000 kilograms) must meet Stage 3/Chapter 3 noise (decibel) level requirements which establish an allowable decibel noise level. Under Annex 16 Stage 4/Chapter 4 requirements, a lower maximum (decibel) level will be mandated, by at least a reduction of 10 decibels from current Stage 3/Chapter 3 levels. Such noise reduction is effected by mixing of the primary hot exhaust gases in an internal mixer with secondary bypass cooling air and by breaking of the single core of exhaust gases into a plurality of smaller cores through use of a first set of lobes positioned internally in the engine.
  • IAO International Civil Aviation Organization
  • a second set of lobes in an external mixer is positioned downstream from the first set at the terminus of the engine.
  • a thrust reverser module is joined to the engine housing or fan case at the engine terminus by use of an attendant mechanism covered by the STANG fairing. Because the engine has specifically designed dimensions, the second set of lobes must be configured to accommodate the existing engine design, which has a terminus exit area dimension of 1,100 square inches (7,097 square centimeters), rather than to reconfigure the engine to fit the second set of lobes. Such engine reconfiguration is impractical and expensive.
  • the second stage or external jet nozzle mixer of the present invention includes a number of lobes, which may be equal in number to those of the first stage or internal mixer, and all of the second stage mixer lobes may be identically formed. As the lobes axially extend outwardly from the mixer attachment to the engine nozzle, they axially inwardly expand from an essentially circular base to an undulating configuration whose apices increase in height.
  • the lobes include complex curvatures whose interior and exterior surfaces greatly enhance mixing respectively of the previously mixed bypass cooling air-hot exhaust gases from the internal mixer and additional ambient cooling air, and thereby also reduce noise.
  • the area encompassed by the lobes remains essentially the same (1,065 to 1,120 sq. inches) as for the jet engine for which it is designed which, for the Pratt & Whitney JT8D-217/219 Series engine, is 1,095 to 1,105 square inches (6,089 to 7,097 square centimeters).
  • the lobe terminus area is consistent with that of the engine in question.
  • the external mixer length is 12 inches ⁇ 3 inches (30.45 cm ⁇ 8 cm).
  • the essentially circular base of the lobes at the mixer inlet has a linear dimension of 39.7 inches (101 centimeters) round, providing an area of 1,223 sq. inches (7,891 square centimeters).
  • the dimension of the mixer circumscribing the lobes at their greatest height is also 39.7 inches (101 centimeters) diameter but, because of the scalloped lobe shape, the area enclosed by the lobes is 1,065 to 1,120 sq. inches (6,089 to 6,403 square centimeters), which matches the area of the existing tailpipe.
  • the exit shape has elliptical shaped lobes and is proportional to a 10 ⁇ 2.5 ellipse (plus or minus 2 inch major axis, and +0.5 inch minor axis). These curve sides help resist distortion caused by the exhaust gas pressure.
  • an important and preferred design parameter is to shape the external mixer of the present invention with a generally cylindrical configuration and with as short a length as possible, so that it does not interfere with the existing thrust reverser doors at the end of the tailpipe.
  • the mixer of the present invention permits the use of existing thrust reversers without necessitating any modification thereto. Only a part of the STANG fairings need to be slightly decreased in their inner dimensions to accommodate the internal mixer. Also, the existing tailpipe is shortened by about 5 inches (12.7 centimeters).
  • the interior surfaces of the lobes force the impinging hot gases, as previously mixed with the secondary bypass cooling air by the first set of lobes of the internal mixer, in all directions towards the interior of the mixer, essentially 45° to 60°, to effect a vigorous mixing of the gases. Simultaneously, additional ambient cooling air is forced from the exterior surfaces of the lobes to mix further with the internally mixed gases. These actions cause the smaller gas cores, which were formed by the first stage mixer, to break into innumerable forms which are both cooler and considerably noise attenuated. In part, the internal contours of the lobes act as flutes to produce a lifting effect which causes the primary hot and cold flows to mix before entering the nozzle.
  • the external contours of the lobes act as chutes which produce a venturi effect and accelerate the cooler secondary flow of ambient air.
  • the lobes thereby act collectively as an injector to force the cooler ambient secondary flow into the previously mixed primary flow as it exits the nozzle. These actions further reduce the noise level.
  • the curve sides of the lobes help resist distortion caused by the exhaust gas pressure.
  • An ameliorative further result is that the accelerated gas/air flow helps to faster move large, previously slowed mixtures to increase the efficiency of the jet engine, by increasing its thrust, that is, an increased thrust specific fuel consumption (TSFC) is estimated to be about a 3% improvement.
  • TSFC thrust specific fuel consumption
  • the axial length of the mixer of the present invention is 12 inches ⁇ 3 inches, which means that there is a lesser distance between the nozzle exit and the buckets of the thrust reverser. The effect of such decreased distance is that more of the thrust from the engine is captured by the buckets and thus utilized to brake the aircraft when needed.
  • the jet nozzle mixer of the present invention fits within and is attachable to the existing engine exit whose area which, as stated above, is 1,095-1,105 square inches (6,261-7,129 square centimeters) exit area for the Pratt & Whitney JT8D-217/219 Series engine.
  • the lobes of the present invention can be made uniform and easily tailored to provide an efficient mixing of the exhaust gases with the ambient air and the attendant reduction in noise. Its uniform dimensions enables its manufacturing costs to be reduced.
  • FIGS. 1 a and 1 b are perspective views of an end portion of a jet engine nozzle assembly to which is attached both a thrust reverser and a second stage external jet nozzle mixer as embodied in the present invention.
  • FIG. 2 is schematic drawing illustrating the interior of the jet engine shown in FIG. 1 with a known first stage internal mixer in the interior of the engine and the second stage inventive external jet nozzle mixer at the terminus of the engine, including the decrease in distance between the jet nozzle mixer of the present invention and the thrust reverser buckets, as compared to its non-use.
  • FIG. 3 a is a view of the engine and its internal mixer shown in FIG. 2 taken along line 3 - 3 thereof
  • FIG. 3 b is a perspective view of the cone and surrounding vanes of the internal mixer.
  • FIG. 4 is a perspective view of the second stage, external mixer assembly of the present invention in which its twelve identically shaped lobes are seen.
  • the four undulating cross-sections, # 1 through # 4 which variously pass through the lobes of the mixer and which extend from the end of the mixer assembly towards its point of attachment to the terminus of the engine, are representative of all planes which pass through all of the lobes.
  • a fifth cross-section # 5 which is circular, extends about the band which anchors the mixer to the engine terminus.
  • a sixth cross-section # 6 is positioned behind the plane of the fifth cross-section # 5 , and is seen in subsequent figures. These six cross-sections are referred to in subsequent figures as defining planes numbered # 1 -# 6 .
  • FIG. 5 is view of the external mixer of the present invention taken perpendicularly to and along the axis of the mixer assembly shown in FIG. 4 , and also depicts how the lobes disperse and break up the hot gas/air mixture.
  • FIGS. 5-1 through 5 - 5 illustrate the areas incorporated by the lobes at their respective cross-sections # 1 -# 5 .
  • the cross-sections, as portrayed or positioned on the interior surfaces of the lobes, define interior mixer areas within their respective planes, respectively of 1,100 square inches (7,097 square centimeters) at plane # 1 ( FIG. 5-1 ), 1,110 square inches (7,162 square centimeters) at plane # 2 ( FIG.
  • FIG. 6 is a side view, taken 90° with respect to the mixer shown in FIG. 5 , of that mixer and its four undulating cross-sections and fifth circular cross-section along planes # 1 -# 5 .
  • the circular configuration of the lobes at plane # 5 extends generally cylindrically with the same general diameter to its end at plane # 6 .
  • FIG. 7 is an enlarged view of a superimposition of the lobes and the same previously illustrated four undulating cross-sections and fifth circular cross-section as shown in FIGS. 4-6 .
  • FIG. 8 is a view of the lobe shown in FIG. 7 looking down upon the apex of the lobe, in which the several cross-sections indicate the varying curvature of the lobe as its extends along the mixer axis through cross-sections or planes # 1 -# 6 .
  • FIG. 9 is a side view of the lobe shown in FIGS. 4-8 and illustrates the several lobe curvatures as it extends along the mixer axis, with specific reference to planes 1 - 6 with its attaching end to the nozzle or tailpipe.
  • FIG. 10 depicts the contour lines of a lobe between its planes # 1 -# 6 , as viewed looking down upon the lobe.
  • FIG. 11 is a view of a specific one of the section curvatures shown in FIG. 9 along with hardware for its attachment to the nozzle or tailpipe.
  • FIG. 12 is a schematic drawing, not to scale, of an engine nozzle assembly and modified STANG fairings for accommodating the jet nozzle mixer embodied in the present invention.
  • FIG. 13 is a perspective view of one of the STANG fairings as modified to accommodate the mixer of the present invention.
  • FIG. 14 is a graph attesting to the improvement in net thrust versus engine pressure ratio in a Pratt & Whitney JT8D-217/219 Series engine when use of the second stage external mixer of the present invention is compared to that of a standard nozzle, in which the engine pressure ratio is defined as the measure of engine exhaust pressure divided by ambient pressure.
  • FIG. 15 is a graph demonstrating the improvement in TSFC (thrust specific fuel consumption) versus thrust in a Pratt & Whitney JT8D-217/219 Series engine when use of the second stage external mixer of the present invention is compared to that of a standard nozzle.
  • TSFC thrust specific fuel consumption
  • FIG. 16 is a graph of preliminary flight test data of a McDonnell-Douglas MD-80 aircraft as evidence of the improvement in fuel consumption in terms of NAMPP (nautical air miles per pound of fuel) versus mach in a Pratt & Whitney JT8D-217/219 Series engine when use of the second stage external mixer of the present invention is compared to that of a standard nozzle.
  • NAMPP nautical air miles per pound of fuel
  • FIGS. 1 a and 1 b illustrate a nozzle assembly 18 relating to, for example, a Pratt & Whitney JT8D-217/219 Series jet engine to which a jet nozzle mixer 20 as embraced by the present invention is attached at its exhaust terminus 19 .
  • Assembly 18 also supports a thrust reverser having a pair of thrust reverser buckets 22 .
  • the attachment of the thrust reverser buckets to assembly 20 is effected by bars 24 which are pivotally linked to a pair of diametrically opposed mechanisms housed within fairings 26 , one of which is shown in FIG. 1 .
  • the fairings are secured at opposite sides of the assembly.
  • the thrust reversers and the linking bars are of conventional design and are unmodified when coupled with the present invention.
  • the fairings are also of conventional design, but a part of the structure covered thereby is slightly modified as will be explained below with respect to FIGS. 12 and 13 .
  • mixer 20 because of its added axial length, is positioned closer to thrust reverser buckets 22 when they are deployed as brakes. Such closer positioning is demonstrated by the different lengths “x” and “y” of FIG. 2 . The ameliorative result of such closer positioning permits the buckets to capture a greater portion of the exhaust for braking purposes than previously obtainable. However, it is important that mixer 20 not be located too close to buckets 22 so that the flow of the redirected exhaust gases are not adversely affected and that the doors, linkages and the mixer are not deleteriously stressed.
  • FIGS. 2, 3 a and 3 b The internal arrangement of nozzle assembly 18 as secured to a jet engine is depicted in FIGS. 2, 3 a and 3 b .
  • An engine 28 includes turbine blades 30 and compressor or fan blades 31 joined together on a common shaft 32 within a two-part housing 34 a and 34 b . For convenience, the burners preceding turbine blades 30 are not shown. Hot exhaust gases exit from the turbine blades as a core 36 .
  • a bypass or fan duct 38 surrounds housing 34 b for affording passage of cooling air, as denoted by arrow-headed lines 39 , from the ambient exterior to first stage or internal jet nozzle mixer 42 of the engine.
  • Core 36 of hot gases is disposed to be mixed with the cooling air within a first stage mixing chamber 40 by use of first stage jet nozzle mixer 42 positioned therein.
  • first stage internal jet nozzle mixer 42 includes two sets of vanes 44 and 46 which are respectively inwardly and outwardly inclined to direct and mix together respectively the cooling air and the hot gases in chamber 40 .
  • Vanes 44 and 46 are positioned around a core terminating in a cone 47 .
  • the total of inwardly directed cooling air vanes 44 and outwardly directed hot gas vanes 46 respectively may number twelve each.
  • This resulting admixture divides core 36 into a smaller cooler central core and twelve surrounding small cores of mixed hot gases and cooling air of different velocities which, nevertheless, are still extremely hot and produce an unacceptably high noise level.
  • These smaller central and surrounding cores pass towards terminus 19 of the nozzle assembly for second stage mixing and cooling by second stage external jet nozzle mixer 20 of the present invention.
  • Mixer 20 may include twelve identical lobes 48 to equal in number the twelve cooling air vanes and the twelve hot gas vanes, and the twelve smaller hot gas cores of the internal mixer.
  • twelve sections, each including a lobe is fabricated and the sections on either side of the lobes are welded together, such as identified by weld lines 50 .
  • the lobes extend from a circular section through a plurality of increasingly undulating portions, such as exemplified by cross-sections # 1 -# 6 .
  • the transition from a round configuration at cross-section # 5 to the scalloped or undulated configuration at cross-section # 1 is a very smooth complex curve and, consequently, minimizes airflow distortion and drag and maximizes the mixing of the hot gases with neighboring air and thereby to reduce noise. This is achieved by using synchronized cross-sections and a plurality of weighted and blending splines between the cross-sections. Such a design is provided using state-of-the-art CAD software.
  • the cross-sections delimit interior mixer areas within the planes defined by the cross-sections, respectively of 1,100 square inches (7,097 square centimeters) at plane # 1 ( FIG. 5-1 ), 1,110 square inches (7,162 square centimeters) at plane # 2 ( FIG. 5-2 ), 1,120 square inches (7,226 square centimeters) at plane # 3 ( FIG. 5-4 ), 1,154 square inches (7,445 square centimeters) at plane # 4 ( FIG. 5-4 ), and 1,223 square inches (7,891 square centimeters) at plane # 5 ( FIG. 5-5 ).
  • the cross-sectional areas from plane # 5 to plane # 1 decreases arithmetically, about 5%, 2.5%, 1.25%, etc.
  • the section extending between cross-sections # 5 and # 6 is an extension from the section adjacent cross-section # 5 and is used to affix mixer 20 to the nozzle terminating the Pratt & Whitney JT8D-217/219 Series engine, and has an equivalent 1,223 square inch (7,891 square centimeter) area.
  • An annular reinforcing support band 52 joins the lobes at their circularly shaped section adjacent cross-section # 5 , while a band ring 54 is joined to lobes 48 at their base sections 55 at their greatest undulation at cross-section # 1 .
  • FIG. 11 also illustrates the attachment of mixer 20 to nozzle assembly or tailpipe 18 .
  • the mixer is secured to terminus 19 of the nozzle assembly and to a doubler ring 70 .
  • Both terminus 19 and the doubler ring are angled outwardly and, compared to prior nozzle assemblies, are shorter by approximately 5 inches.
  • the interior surfaces of the lobes force the impinging hot gas-bypass cooling air mixture from internal mixer 42 in all directions towards the interior of internal mixer 20 , that is, essentially 45° to 60° as illustrated by multiple arrow-headed lines 56 in FIG. 5 , to effect a vigorous mixing of the gases.
  • additional ambient cooling air is forced from the exterior surfaces of the lobes to mix further with the internally mixed gases.
  • This lifting effect causes the primary hot and cold flows to mix before entering the nozzle.
  • the external contours of the lobes which act as chutes 66 , are designed to act as a multitude of venturis, and thus to accelerate the cooler secondary flow of ambient air.
  • This arrangement effectively forms an injector to force the cooler ambient secondary flow into the previously mixed primary flow as it exits the nozzle. This action further reduces the noise level.
  • dimples 72 are formed on both sides of band 54 of the external mixer and act as vortex generators to prevent the mixed gas flow from attaching to band 54 and thereby to enhance the mixing action.
  • This afore-mentioned acceleration also helps to increase the efficiency of the fuel-air burning in the engine.
  • By producing an increased flow the exhaust gases are more rapidly exhausted from the engine and thereby the need for the engine and its bypass compressor to expend energy in moving these gases is alleviated.
  • the lobes are elliptically shaped, being proportional to a 10 ⁇ 2.5 ellipse, plus or minus 2 inches (5 centimeters) major axis, and plus or minus 0.5 inch (1.3 centimeter) minor axis. These curved sides help resist distortion caused by the exhaust gas pressure.
  • mixer 20 such as illustrated in FIGS. 4, 5 , et seq., has a 1,065 sq. inch to 1,100 square inch (6,089 to 7,097 square centimeters) area encompassed by the lobes at plane # 1 and a 1,223 square inch (7,891 square centimeters) area at plane # 5 , where the mixer is joined to nozzle assembly 18 , it is possible to use the mixer without any modification of thrust reversers 22 . As a result, it is necessary only to slightly reconfigure the structure covered by fairings 26 . Such reconfiguration is depicted in FIGS. 12 and 13 , and is effected by removing only a small portion from each of such structure, specifically that portion indicated by parallel dashed lines 58 . Further, a tongue 59 is also removed.
  • the mixer has as short a length as is possible, e.g., 12 inches ⁇ 3 inches (30.45 cm ⁇ 8 cm).
  • the lobe shape starts with a circular or rounded configuration at 39.7 inches (101 centimeters) and terminates with a scalloped or undulated configuration at the same diameter (39.7 inches or 101 centimeters) and an area of 1,065 sq. inches to 1,100 sq. inches (6,089 to 7,097 square centimeters), which matches the existing tailpipe area.
  • the mixer By keeping the mixer short, it will not interfere with the existing thrust reverser doors at the end of the tailpipe.
  • the mixer is designed so that it can be attached to the existing tailpipe with minimum impact on exiting components, such as the thrust reverser, thrust reverser doors, Stang fairings, outer fairings.
  • the mixer has elliptically shaped lobes whose shapes are proportional to a 10 ⁇ 2.5 ellipse (plus or minus 2 inch major axis, and plus or minus 0.5 inch minor axis). These curved sides help to resist distortion caused by exhaust gas pressure.
  • the mixer inner lobe surfaces ramps the exhaust gases inward and, at the same time, the outer surface draws outside air into the mixer using a type of NACA duct (airfoil air scoop) so that, when the hot gases and the cooling air is mixed, the exhaust noise is reduced.
  • NACA duct airfoil air scoop
  • the contour lines of the lobed surfaces form a uniform initial slope, which is desirable to ensure even pressure as the exhaust gases are redirected inward.
  • I. Twelve lobes may be to match the existing twelve vanes in the engine that swirl and spin the exhaust gases as they leave the engine.
  • the twelve “hot spots” inside the tailpipe, which are produced by the existing vanes, are broken up by, for example, the twelve lobes of the present invention, thereby minimizing any undesirable hot spots.
  • the lobe shape forms a complex compound surface, with as large as possible employ of radii used at all locations so as to minimize drag and to allow for the smoothest possible gas flow redirection.
  • FIG. 14 discloses that, based upon a reasonable match for all engine parameters, such as engine revolutions per minute (rpm), exhaust gas temperature (EGT) and fuel pump data, the present invention demonstrates an increase in thrust at the mid range of engine pressure ratio (EPR), that is, engine exhaust pressure divided by ambient pressure.
  • EPR engine pressure ratio
  • FIG. 15 reveals that the present invention, within a mid thrust range of 7,000 to 15,000 pounds of thrust, improves upon the TFC (specific fuel consumption) by a factor of approximately 2% to 3%.
  • TFC specific fuel consumption
  • FIG. 15 reveals that the present invention, within a mid thrust range of 7,000 to 15,000 pounds of thrust, improves upon the TFC (specific fuel consumption) by a factor of approximately 2% to 3%.
  • An engine average fuel burn of 7,000 pounds of fuel per hour converts into an approximate consumption of 1,000 gallons per hour of fuel.
  • FIG. 16 compares the improvement in nautical air miles per pound of fuel (NAMPP) versus mach number for a McDonnell-Douglas MD-80 aircraft through use vis-a-vis non-use of the present invention.
  • NAMPP nautical air miles per pound of fuel
  • preliminary flight data shows an increased NAMPP of the “JET nozzle” over all points on the curve when employing the present invention over its non-use “baseline nozzle.”

Abstract

An external jet nozzle mixer includes identically formed lobes. The external mixer works with the internal mixer further to mix the engine internal bypass flow with the internal jet engine core flow to level the disparate flow velocities, to reduce the peak velocities from the jet engine core and increase the lower bypass velocities of the engine internal bypass flow, and thereby reduce noise. The internal lobe contours act as lifting flutes, causing mixing of the primary hot and cold flows to mix before exiting the nozzle. The external lobe contours act as venturi chutes, accelerating the cooler ambient secondary air flow. The lobes thus act collectively as an injector to force the cooler ambient secondary flow into the previously mixed primary flow as it exits the nozzle. Also obtained is an increased thrust efficiency and a consequently decreased fuel consumption and engine emissions.

Description

    CROSS-REFERENCE TO RELATED APPLICATIONS
  • This application is a continuation of application Ser. No. 10/314,101, filed Dec. 7, 2002, which claims the benefit of U.S. Provisional Application No. 60/340,369, filed Dec. 7, 2001, each of which is incorporated herein by reference thereto, respectively.
  • FIELD OF THE INVENTION
  • The present invention relates to jet nozzle mixers for aircraft jet engines and, in particular, to improvements in effecting a greater cooling and a lower noise level in exhaust gases emanating from such engines and in increasing power and fuel efficiency.
  • BACKGROUND
  • Noise (decibel) level in jet aircraft engines is established by laws and regulations, specifically promulgated by the International Civil Aviation Organization (ICAO), Annex 16. At present, commercial jet aircraft weighing over 75,000 pounds (34,000 kilograms) must meet Stage 3/Chapter 3 noise (decibel) level requirements which establish an allowable decibel noise level. Under Annex 16 Stage 4/Chapter 4 requirements, a lower maximum (decibel) level will be mandated, by at least a reduction of 10 decibels from current Stage 3/Chapter 3 levels. Such noise reduction is effected by mixing of the primary hot exhaust gases in an internal mixer with secondary bypass cooling air and by breaking of the single core of exhaust gases into a plurality of smaller cores through use of a first set of lobes positioned internally in the engine. For some engines, a second set of lobes in an external mixer is positioned downstream from the first set at the terminus of the engine. A thrust reverser module is joined to the engine housing or fan case at the engine terminus by use of an attendant mechanism covered by the STANG fairing. Because the engine has specifically designed dimensions, the second set of lobes must be configured to accommodate the existing engine design, which has a terminus exit area dimension of 1,100 square inches (7,097 square centimeters), rather than to reconfigure the engine to fit the second set of lobes. Such engine reconfiguration is impractical and expensive. Therefore, the direction towards meeting Stage 3/Chapter 3 noise requirements has been involved in developing a variously configured second set of lobes whose design does not always meet such requirements and, when the lobe design does, the lobes are difficult and expensive to manufacture and the mixer is expensive to be retrofitted to the engine.
  • Some engines have not employed the use of a second set of lobes or an external mixer, specifically one produced by Pratt & Whitney, in their JT8D-217/219 Series. Currently, this engine includes an internal 12 lobe mixer and is only certified to Stage 3/Chapter 3 noise levels. There has been a desire to qualify this particular engine to Stage 4/Chapter 4 noise levels, but to minimize the costs of doing so with, preferably, no changes in its thrust reversal components primarily because of cost and other economic reasons. To bring this engine to Stage 4/Chapter 4 noise levels, an additional 2 decibel reduction in jet noise is required. Such an upgrading is a challenge that has not been met.
  • BRIEF SUMMARY OF THE INVENTION
  • These and other problems are avoided and the Stage 4/Chapter 4 requirements are both met and surpassed by the present invention, not only for the above-mentioned Pratt & Whitney JT8D-217/219 Series engine but also for other engines. The second stage or external jet nozzle mixer of the present invention includes a number of lobes, which may be equal in number to those of the first stage or internal mixer, and all of the second stage mixer lobes may be identically formed. As the lobes axially extend outwardly from the mixer attachment to the engine nozzle, they axially inwardly expand from an essentially circular base to an undulating configuration whose apices increase in height. The lobes include complex curvatures whose interior and exterior surfaces greatly enhance mixing respectively of the previously mixed bypass cooling air-hot exhaust gases from the internal mixer and additional ambient cooling air, and thereby also reduce noise. At their terminus, the area encompassed by the lobes remains essentially the same (1,065 to 1,120 sq. inches) as for the jet engine for which it is designed which, for the Pratt & Whitney JT8D-217/219 Series engine, is 1,095 to 1,105 square inches (6,089 to 7,097 square centimeters). For other engines, the lobe terminus area is consistent with that of the engine in question.
  • For the Pratt & Whitney JT8D-217/219 Series engine, for example, the external mixer length is 12 inches±3 inches (30.45 cm±8 cm). The essentially circular base of the lobes at the mixer inlet has a linear dimension of 39.7 inches (101 centimeters) round, providing an area of 1,223 sq. inches (7,891 square centimeters). At the mixer outlet at the full height of the regularly undulating lobes, the dimension of the mixer circumscribing the lobes at their greatest height is also 39.7 inches (101 centimeters) diameter but, because of the scalloped lobe shape, the area enclosed by the lobes is 1,065 to 1,120 sq. inches (6,089 to 6,403 square centimeters), which matches the area of the existing tailpipe.
  • The exit shape has elliptical shaped lobes and is proportional to a 10×2.5 ellipse (plus or minus 2 inch major axis, and +0.5 inch minor axis). These curve sides help resist distortion caused by the exhaust gas pressure.
  • Consistent with the above discussion, an important and preferred design parameter is to shape the external mixer of the present invention with a generally cylindrical configuration and with as short a length as possible, so that it does not interfere with the existing thrust reverser doors at the end of the tailpipe. As a result, the mixer of the present invention permits the use of existing thrust reversers without necessitating any modification thereto. Only a part of the STANG fairings need to be slightly decreased in their inner dimensions to accommodate the internal mixer. Also, the existing tailpipe is shortened by about 5 inches (12.7 centimeters).
  • Functionally, the interior surfaces of the lobes force the impinging hot gases, as previously mixed with the secondary bypass cooling air by the first set of lobes of the internal mixer, in all directions towards the interior of the mixer, essentially 45° to 60°, to effect a vigorous mixing of the gases. Simultaneously, additional ambient cooling air is forced from the exterior surfaces of the lobes to mix further with the internally mixed gases. These actions cause the smaller gas cores, which were formed by the first stage mixer, to break into innumerable forms which are both cooler and considerably noise attenuated. In part, the internal contours of the lobes act as flutes to produce a lifting effect which causes the primary hot and cold flows to mix before entering the nozzle. The external contours of the lobes act as chutes which produce a venturi effect and accelerate the cooler secondary flow of ambient air. The lobes thereby act collectively as an injector to force the cooler ambient secondary flow into the previously mixed primary flow as it exits the nozzle. These actions further reduce the noise level. Further, the curve sides of the lobes help resist distortion caused by the exhaust gas pressure. An ameliorative further result is that the accelerated gas/air flow helps to faster move large, previously slowed mixtures to increase the efficiency of the jet engine, by increasing its thrust, that is, an increased thrust specific fuel consumption (TSFC) is estimated to be about a 3% improvement. Such increased TSFC occurs through better dynamic mixing of the bypass or fan duct and turbine exhaust gases. It addresses the problem of the transfer from a hot, high velocity volume to a cooler, slower velocity volume. This mixing levels the disparate flow velocities attendant with the jet engine exhaust, reduces the peak velocities from the jet engine core and increases the lower bypass velocities of the jet engine internal bypass flow. Because noise is a function of jet exhaust velocity to the 7th power, and because peak velocities from the core flow are reduced, the jet noise is thereby reduced.
  • As stated above, the axial length of the mixer of the present invention is 12 inches±3 inches, which means that there is a lesser distance between the nozzle exit and the buckets of the thrust reverser. The effect of such decreased distance is that more of the thrust from the engine is captured by the buckets and thus utilized to brake the aircraft when needed.
  • Several advantages are derived from this arrangement. The jet nozzle mixer of the present invention fits within and is attachable to the existing engine exit whose area which, as stated above, is 1,095-1,105 square inches (6,261-7,129 square centimeters) exit area for the Pratt & Whitney JT8D-217/219 Series engine. The lobes of the present invention can be made uniform and easily tailored to provide an efficient mixing of the exhaust gases with the ambient air and the attendant reduction in noise. Its uniform dimensions enables its manufacturing costs to be reduced. The need to modify the existing thrust reverser per se is avoided because the mixer is fittable and attachable to the existing engine exit; only minor dimensional changes in the existing STANG fairing, and tailpipe and outer barrel are required without otherwise needing any change in other components such as the thrust reverser, the thrust reverser doors, and their linkages. Efficiency in jet engine operation is increased, with concomitant saving of fuel and costs thereof. Thrust reverser braking of the aircraft is improved.
  • Other aims and advantages, as well as a more complete understanding of the present invention, will appear from the following explanation of an exemplary embodiment and the accompanying drawings thereof.
  • BRIEF DESCRIPTION OF THE DRAWINGS
  • FIGS. 1 a and 1 b are perspective views of an end portion of a jet engine nozzle assembly to which is attached both a thrust reverser and a second stage external jet nozzle mixer as embodied in the present invention.
  • FIG. 2 is schematic drawing illustrating the interior of the jet engine shown in FIG. 1 with a known first stage internal mixer in the interior of the engine and the second stage inventive external jet nozzle mixer at the terminus of the engine, including the decrease in distance between the jet nozzle mixer of the present invention and the thrust reverser buckets, as compared to its non-use.
  • FIG. 3 a is a view of the engine and its internal mixer shown in FIG. 2 taken along line 3-3 thereof, and FIG. 3 b is a perspective view of the cone and surrounding vanes of the internal mixer.
  • FIG. 4 is a perspective view of the second stage, external mixer assembly of the present invention in which its twelve identically shaped lobes are seen. The four undulating cross-sections, #1 through #4, which variously pass through the lobes of the mixer and which extend from the end of the mixer assembly towards its point of attachment to the terminus of the engine, are representative of all planes which pass through all of the lobes. A fifth cross-section # 5, which is circular, extends about the band which anchors the mixer to the engine terminus. A sixth cross-section # 6 is positioned behind the plane of the fifth cross-section # 5, and is seen in subsequent figures. These six cross-sections are referred to in subsequent figures as defining planes numbered #1-#6.
  • FIG. 5 is view of the external mixer of the present invention taken perpendicularly to and along the axis of the mixer assembly shown in FIG. 4, and also depicts how the lobes disperse and break up the hot gas/air mixture. FIGS. 5-1 through 5-5 illustrate the areas incorporated by the lobes at their respective cross-sections #1-#5. The cross-sections, as portrayed or positioned on the interior surfaces of the lobes, define interior mixer areas within their respective planes, respectively of 1,100 square inches (7,097 square centimeters) at plane #1 (FIG. 5-1), 1,110 square inches (7,162 square centimeters) at plane #2 (FIG. 5-2), 1,120 square inches (7,226 square centimeters) at plane #3 (FIG. 5-3), 1,154 square inches (7,445 square centimeters) at plane #4 (FIG. 5-4), and 1,223 square inches (7,891 square centimeters) at plane #5 (FIG. 5-5) which extends into plane 6 for attachment to the existing Pratt & Whitney JT8D-217/219 Series engine.
  • FIG. 6 is a side view, taken 90° with respect to the mixer shown in FIG. 5, of that mixer and its four undulating cross-sections and fifth circular cross-section along planes #1-#5. The circular configuration of the lobes at plane # 5 extends generally cylindrically with the same general diameter to its end at plane # 6.
  • FIG. 7 is an enlarged view of a superimposition of the lobes and the same previously illustrated four undulating cross-sections and fifth circular cross-section as shown in FIGS. 4-6.
  • FIG. 8 is a view of the lobe shown in FIG. 7 looking down upon the apex of the lobe, in which the several cross-sections indicate the varying curvature of the lobe as its extends along the mixer axis through cross-sections or planes #1-#6.
  • FIG. 9 is a side view of the lobe shown in FIGS. 4-8 and illustrates the several lobe curvatures as it extends along the mixer axis, with specific reference to planes 1-6 with its attaching end to the nozzle or tailpipe.
  • FIG. 10 depicts the contour lines of a lobe between its planes #1-#6, as viewed looking down upon the lobe.
  • FIG. 11 is a view of a specific one of the section curvatures shown in FIG. 9 along with hardware for its attachment to the nozzle or tailpipe.
  • FIG. 12 is a schematic drawing, not to scale, of an engine nozzle assembly and modified STANG fairings for accommodating the jet nozzle mixer embodied in the present invention.
  • FIG. 13 is a perspective view of one of the STANG fairings as modified to accommodate the mixer of the present invention.
  • FIG. 14 is a graph attesting to the improvement in net thrust versus engine pressure ratio in a Pratt & Whitney JT8D-217/219 Series engine when use of the second stage external mixer of the present invention is compared to that of a standard nozzle, in which the engine pressure ratio is defined as the measure of engine exhaust pressure divided by ambient pressure.
  • FIG. 15 is a graph demonstrating the improvement in TSFC (thrust specific fuel consumption) versus thrust in a Pratt & Whitney JT8D-217/219 Series engine when use of the second stage external mixer of the present invention is compared to that of a standard nozzle.
  • FIG. 16 is a graph of preliminary flight test data of a McDonnell-Douglas MD-80 aircraft as evidence of the improvement in fuel consumption in terms of NAMPP (nautical air miles per pound of fuel) versus mach in a Pratt & Whitney JT8D-217/219 Series engine when use of the second stage external mixer of the present invention is compared to that of a standard nozzle.
  • DETAILED DESCRIPTION OF THE INVENTION
  • Because the present invention was devised particularly with respect to the Pratt & Whitney JT8D-217/219 Series engine, the following discussion will be directed specifically thereto; however, it is to be understood that the present invention is equally relevant for use in other jet engines and, therefore, is not to be limited to a specific jet engine.
  • Accordingly, FIGS. 1 a and 1 b illustrate a nozzle assembly 18 relating to, for example, a Pratt & Whitney JT8D-217/219 Series jet engine to which a jet nozzle mixer 20 as embraced by the present invention is attached at its exhaust terminus 19. Assembly 18 also supports a thrust reverser having a pair of thrust reverser buckets 22. The attachment of the thrust reverser buckets to assembly 20 is effected by bars 24 which are pivotally linked to a pair of diametrically opposed mechanisms housed within fairings 26, one of which is shown in FIG. 1. The fairings are secured at opposite sides of the assembly. The thrust reversers and the linking bars are of conventional design and are unmodified when coupled with the present invention. The fairings are also of conventional design, but a part of the structure covered thereby is slightly modified as will be explained below with respect to FIGS. 12 and 13.
  • As shown also in FIG. 2, mixer 20, because of its added axial length, is positioned closer to thrust reverser buckets 22 when they are deployed as brakes. Such closer positioning is demonstrated by the different lengths “x” and “y” of FIG. 2. The ameliorative result of such closer positioning permits the buckets to capture a greater portion of the exhaust for braking purposes than previously obtainable. However, it is important that mixer 20 not be located too close to buckets 22 so that the flow of the redirected exhaust gases are not adversely affected and that the doors, linkages and the mixer are not deleteriously stressed.
  • The internal arrangement of nozzle assembly 18 as secured to a jet engine is depicted in FIGS. 2, 3 a and 3 b. An engine 28 includes turbine blades 30 and compressor or fan blades 31 joined together on a common shaft 32 within a two- part housing 34 a and 34 b. For convenience, the burners preceding turbine blades 30 are not shown. Hot exhaust gases exit from the turbine blades as a core 36. A bypass or fan duct 38 surrounds housing 34 b for affording passage of cooling air, as denoted by arrow-headed lines 39, from the ambient exterior to first stage or internal jet nozzle mixer 42 of the engine. Core 36 of hot gases is disposed to be mixed with the cooling air within a first stage mixing chamber 40 by use of first stage jet nozzle mixer 42 positioned therein. As best seen in FIGS. 2, 3 a and 3 b, first stage internal jet nozzle mixer 42 includes two sets of vanes 44 and 46 which are respectively inwardly and outwardly inclined to direct and mix together respectively the cooling air and the hot gases in chamber 40. Vanes 44 and 46 are positioned around a core terminating in a cone 47. As stated above, consistent with the Pratt & Whitney JT8D-217/219 Series jet engine design, the total of inwardly directed cooling air vanes 44 and outwardly directed hot gas vanes 46 respectively may number twelve each. This resulting admixture divides core 36 into a smaller cooler central core and twelve surrounding small cores of mixed hot gases and cooling air of different velocities which, nevertheless, are still extremely hot and produce an unacceptably high noise level. These smaller central and surrounding cores pass towards terminus 19 of the nozzle assembly for second stage mixing and cooling by second stage external jet nozzle mixer 20 of the present invention.
  • Second stage external jet nozzle mixer 20 and its component parts is illustrated in FIGS. 4-11. Mixer 20, as illustrated, may include twelve identical lobes 48 to equal in number the twelve cooling air vanes and the twelve hot gas vanes, and the twelve smaller hot gas cores of the internal mixer. For ease of manufacture, twelve sections, each including a lobe, is fabricated and the sections on either side of the lobes are welded together, such as identified by weld lines 50. Combined, the lobes extend from a circular section through a plurality of increasingly undulating portions, such as exemplified by cross-sections #1-#6. The transition from a round configuration at cross-section # 5 to the scalloped or undulated configuration at cross-section # 1 is a very smooth complex curve and, consequently, minimizes airflow distortion and drag and maximizes the mixing of the hot gases with neighboring air and thereby to reduce noise. This is achieved by using synchronized cross-sections and a plurality of weighted and blending splines between the cross-sections. Such a design is provided using state-of-the-art CAD software.
  • As stated above, the cross-sections, as portrayed on the interior surfaces of the lobes and depicted by shading in FIGS. 5-1 through 5-5, delimit interior mixer areas within the planes defined by the cross-sections, respectively of 1,100 square inches (7,097 square centimeters) at plane #1 (FIG. 5-1), 1,110 square inches (7,162 square centimeters) at plane #2 (FIG. 5-2), 1,120 square inches (7,226 square centimeters) at plane #3 (FIG. 5-4), 1,154 square inches (7,445 square centimeters) at plane #4 (FIG. 5-4), and 1,223 square inches (7,891 square centimeters) at plane #5 (FIG. 5-5). The cross-sectional areas from plane # 5 to plane # 1 decreases arithmetically, about 5%, 2.5%, 1.25%, etc.
  • The section extending between cross-sections # 5 and #6 is an extension from the section adjacent cross-section # 5 and is used to affix mixer 20 to the nozzle terminating the Pratt & Whitney JT8D-217/219 Series engine, and has an equivalent 1,223 square inch (7,891 square centimeter) area. An annular reinforcing support band 52 (see particularly FIG. 11) joins the lobes at their circularly shaped section adjacent cross-section # 5, while a band ring 54 is joined to lobes 48 at their base sections 55 at their greatest undulation at cross-section # 1.
  • FIG. 11 also illustrates the attachment of mixer 20 to nozzle assembly or tailpipe 18. Specifically, the mixer is secured to terminus 19 of the nozzle assembly and to a doubler ring 70. Both terminus 19 and the doubler ring are angled outwardly and, compared to prior nozzle assemblies, are shorter by approximately 5 inches.
  • As shown, for example in FIGS. 7 and 11, the interior surfaces of the lobes force the impinging hot gas-bypass cooling air mixture from internal mixer 42 in all directions towards the interior of internal mixer 20, that is, essentially 45° to 60° as illustrated by multiple arrow-headed lines 56 in FIG. 5, to effect a vigorous mixing of the gases. At the same time, additional ambient cooling air is forced from the exterior surfaces of the lobes to mix further with the internally mixed gases. These actions cause the smaller gas cores from internal mixer 42 to break into myriad forms which are both cooler and considerably noise attenuated. In part, the internal contours of the lobes act as flutes or channels 64 to produce a similar aerodynamic action as the skins of the airplane wings to produce a lifting effect. This lifting effect causes the primary hot and cold flows to mix before entering the nozzle. The external contours of the lobes, which act as chutes 66, are designed to act as a multitude of venturis, and thus to accelerate the cooler secondary flow of ambient air. This arrangement effectively forms an injector to force the cooler ambient secondary flow into the previously mixed primary flow as it exits the nozzle. This action further reduces the noise level.
  • In addition, dimples 72 are formed on both sides of band 54 of the external mixer and act as vortex generators to prevent the mixed gas flow from attaching to band 54 and thereby to enhance the mixing action.
  • This afore-mentioned acceleration also helps to increase the efficiency of the fuel-air burning in the engine. By producing an increased flow, the exhaust gases are more rapidly exhausted from the engine and thereby the need for the engine and its bypass compressor to expend energy in moving these gases is alleviated.
  • In addition, the lobes are elliptically shaped, being proportional to a 10×2.5 ellipse, plus or minus 2 inches (5 centimeters) major axis, and plus or minus 0.5 inch (1.3 centimeter) minor axis. These curved sides help resist distortion caused by the exhaust gas pressure.
  • Because mixer 20, such as illustrated in FIGS. 4, 5, et seq., has a 1,065 sq. inch to 1,100 square inch (6,089 to 7,097 square centimeters) area encompassed by the lobes at plane # 1 and a 1,223 square inch (7,891 square centimeters) area at plane # 5, where the mixer is joined to nozzle assembly 18, it is possible to use the mixer without any modification of thrust reversers 22. As a result, it is necessary only to slightly reconfigure the structure covered by fairings 26. Such reconfiguration is depicted in FIGS. 12 and 13, and is effected by removing only a small portion from each of such structure, specifically that portion indicated by parallel dashed lines 58. Further, a tongue 59 is also removed.
  • The following points, although not exclusive, may be advanced in summary of the illustrated embodiments of the present invention.
  • A. As an important design parameter, the mixer has as short a length as is possible, e.g., 12 inches±3 inches (30.45 cm±8 cm). The lobe shape starts with a circular or rounded configuration at 39.7 inches (101 centimeters) and terminates with a scalloped or undulated configuration at the same diameter (39.7 inches or 101 centimeters) and an area of 1,065 sq. inches to 1,100 sq. inches (6,089 to 7,097 square centimeters), which matches the existing tailpipe area. By keeping the mixer short, it will not interfere with the existing thrust reverser doors at the end of the tailpipe.
  • B. The mixer is designed so that it can be attached to the existing tailpipe with minimum impact on exiting components, such as the thrust reverser, thrust reverser doors, Stang fairings, outer fairings.
  • C. The mixer has elliptically shaped lobes whose shapes are proportional to a 10×2.5 ellipse (plus or minus 2 inch major axis, and plus or minus 0.5 inch minor axis). These curved sides help to resist distortion caused by exhaust gas pressure.
  • D. The transition in the lobes from a round to a scalloped shape forms a very smooth curve in order to minimize airflow distortion and drag and to maximize the mixing of the hot gases with neighboring air. This is achieved by using six synchronized cross-sections and many weighted and blending splines between the cross-sections. The design was achieved using state-of-the-art CAD software, Surfcam, from Surfware, Inc.
  • E. The cross-sectional area of the mixer, taken along its axis, decreases arithmetically, about 5%, 2.5%, 1.25%, etc., until its terminus is reached.
  • F. Rather than simply splitting the air flow, the mixer inner lobe surfaces ramps the exhaust gases inward and, at the same time, the outer surface draws outside air into the mixer using a type of NACA duct (airfoil air scoop) so that, when the hot gases and the cooling air is mixed, the exhaust noise is reduced.
  • G. The contour lines of the lobed surfaces form a uniform initial slope, which is desirable to ensure even pressure as the exhaust gases are redirected inward.
  • H. Testing of the final lobe shape design with models ensured that the lobes would be formed with relative ease from a flat sheet, and with minimum distortion or strain which would be otherwise caused by material stretching and compressing as the flat sheet is forced into the desired configuration. Such ease of formation is amenable to selection of the preferred material which comprises an aerospace alloy, Inconel 625, a difficult material to work.
  • I. Twelve lobes may be to match the existing twelve vanes in the engine that swirl and spin the exhaust gases as they leave the engine. The twelve “hot spots” inside the tailpipe, which are produced by the existing vanes, are broken up by, for example, the twelve lobes of the present invention, thereby minimizing any undesirable hot spots.
  • J. The lobe shape forms a complex compound surface, with as large as possible employ of radii used at all locations so as to minimize drag and to allow for the smoothest possible gas flow redirection.
  • Preliminary testing of the present invention, as used in a Pratt & Whitney JT8D-217/219 Series jet engine, has disclosed decided improvements in performance as compared to conventional technology. Such data, as shown in FIGS. 14-16, are based upon present testing. It is therefore to be understood that final test results may evidence different data. Notwithstanding, as shown in these graphical representations of preliminary test data, the external or second stage mixer of the present invention demonstrates improved performance over that obtainable with conventional systems.
  • FIG. 14 discloses that, based upon a reasonable match for all engine parameters, such as engine revolutions per minute (rpm), exhaust gas temperature (EGT) and fuel pump data, the present invention demonstrates an increase in thrust at the mid range of engine pressure ratio (EPR), that is, engine exhaust pressure divided by ambient pressure. These tests were conducted by use of the external or second stage mixer of the present invention as compared to use of a standard nozzle (Serial Number 48099 as detailed in a United Technologies Corporation (UTC) document for its Pratt & Whitney engines, entitled “JT8D-209, -217, -217A, -217C, -219, TURBOFAN ENGINES ENGINE MANUAL PART NO.773128” bearing an initial issue date of Jul. 1, 1979 and revised Nov. 15, 2001.
  • FIG. 15 reveals that the present invention, within a mid thrust range of 7,000 to 15,000 pounds of thrust, improves upon the TFC (specific fuel consumption) by a factor of approximately 2% to 3%. The following example is given to demonstrate the economic benefits obtained by assuming a 2% increase in fuel consumption. An engine average fuel burn of 7,000 pounds of fuel per hour converts into an approximate consumption of 1,000 gallons per hour of fuel. Based upon an assumed yearly flight usage of a McDonnell-Douglas MD-80 aircraft of about 2,000 hours per year, the aircraft consumes about 2,000,000 gallons of fuel per year. At a cost of $1.00 per gallon, the annual fuel cost for such an aircraft would be $2,000,000. Therefore, for a 2% improvement in fuel consumption as provided by the present invention, the saving would amount to $40,000 per aircraft.
  • FIG. 16 compares the improvement in nautical air miles per pound of fuel (NAMPP) versus mach number for a McDonnell-Douglas MD-80 aircraft through use vis-a-vis non-use of the present invention. Here, preliminary flight data shows an increased NAMPP of the “JET nozzle” over all points on the curve when employing the present invention over its non-use “baseline nozzle.”
  • It is to be understood that, in the foregoing exposition where dimensions, areas, etc., are expressed in English system units and, parenthetically, in metric system units, the English unit system shall take precedence in the event of any error in conversion from the English unit system to the metric unit system.
  • Although the invention has been described with respect to a particular embodiment thereof, it should be realized that various changes and modifications may be made therein without departing from the spirit and scope of the invention.

Claims (22)

1. In a nozzle for a jet engine enclosed within a tubular housing and also having an existing first stage mixer within said housing which includes a number of first stage lobes, and having a thrust reverser having a stowed position fixed relative to the tubular housing and having a deployed position, the improvement comprising a second stage mixer coupled to the exhaust aperture of said tubular housing to receive gas exhausted from the first stage mixer, said second stage mixer having a plurality of identically formed lobes;
the second stage mixer being unenclosed by a housing and having an exterior in contact with the ambient air to allow the air to enter the mixer and mix with the gas from the first stage mixer;
the second stage mixer having a trailing edge forming the exit terminus for the gas flow; and
said thrust reverser in the deployed position being positioned downstream of the exit terminus to reverse the gas flow.
2. The improvement according to claim 1 wherein the existing first stage lobes mix hot gases from combustion in the jet engine with bypass air flowing through said tubular housing to diminish the velocity of the hot gases, and in which said second stage lobes have interior surfaces which are configured to further mix the mixed hot gas-bypass air mixture and further to level the velocities of the mixed hot gas-bypass air mixture and thereby reduce jet engine noise.
3. The improvement according to claim 2 in which said second stage lobes have exterior surfaces which are configured to draw ambient air external to said tubular housing into mixing with the further mixed hot gas-bypass air mixture and to still further level the velocities of the mixed hot gas-bypass air mixture and thereby further reduce the jet engine noise.
4. The improvement according to claim 1 in which said second stage identically formed lobes are configured (a) on their interior surfaces to force impinging hot gases exiting at said exhaust aperture of said tubular housing surrounding said jet engine towards the interior of said second stage mixer and (b) on their exterior surfaces to mix with ambient cooling air exterior to said tubular housing to form gas cores which are smaller than those formed by the first stage mixer and to break the smaller gas cores into innumerable forms which are both cooler and noise attenuated.
5. The improvement according to claim 4 in which said interior surfaces of said second stage lobes are shaped to accelerate flow of the hot gas-cooling air mixture and thereby to increase the efficiency of the jet engine thrust.
6. The improvement according to claim 4 in which said interior and exterior surfaces of said second stage lobes act respectively as flutes and chutes.
7. The improvement according to claim 1 in which said second stage mixer lobes increase in height from a circular configuration adjacent to the first stage mixer to an undulating configuration at the terminus of said second stage mixer.
8. The improvement according to claim 7 wherein the jet engine has a dimensionally existing terminus exit area defined by said exhaust aperture of said tubular housing, and in which said second stage mixer lobes at the second stage mixer terminus have an area whose dimension is generally the same as that of the jet engine terminus exit area.
9. The improvement according to claim 8 wherein the jet engine terminus exit area and the second stage mixer terminus area are both 1,065 to 1120 square inches.
10. The improvement according to claim 8 wherein a preexisting thrust reverser including conventional thrust reverser doors are secured to a mount rearward of said exhaust aperture, and in which said second stage mixer has an axial length which is sufficiently short as to not interfere with movement of the thrust reverser doors.
11. The improvement according to claim 10 wherein the preexisting thrust reverser and its thrust reverser doors are secured rearward of said exhaust aperture by mechanisms enclosed by STANG fairings, and in which a part of said STANG fairing is slightly decreased in its inner dimension to accommodate said second stage mixer.
12. The improvement according to claim 8 in which said second stage lobes are configured (a) on their interior surfaces to force impinging hot gases from the jet engine towards the interior of said second stage mixer and (b) on their exterior surfaces to mix with ambient cooling air exterior of said tubular housing, to form gas cores which are smaller than those formed by the first stage mixer and to break the smaller gas cores into innumerable forms which are both cooler and noise attenuated.
13. The improvement according to claim 12 in which said second stage lobes are elliptically shaped to help resist distortion otherwise caused by pressure of exhaust gasses.
14. The improvement according to claim 7 in which the increase in height of said second stage mixer lobes from their circular configuration to their undulating configuration forms a smooth curve transition to minimize airflow distortion and drag and to maximize mixing of hot gases from the engine with ambient air.
15. The improvement according to claim 1 in which said second stage mixer is formed from a plurality of secured-together sections, each including one of said second stage lobes.
16. The improvement according to claim 1 in which said second stage lobes are configured to employ the largest possible radii for all surfaces thereof to minimize drag and allow for a smoothest possible redirection of gas flow therealong.
17. The improvement according to claim 1 wherein a thrust reverser having buckets which are deployable into a thrust-reversing position is operatively mounted rearward from said exhaust aperture, and wherein said second stage lobes have an axial length which decreases the distance otherwise existing between the exhaust aperture and the thrust reverser buckets when deployed into their thrust-reversing position for enabling increased capture of the thrust by the buckets.
18. The improvement according to claim 2 further including dimples coupled to the terminus of said second stage for acting as vortex generators to prevent the mixed gas flow from attaching to said lobe terminus and thereby to enhance the mixing action.
19. In a nozzle for a jet engine enclosed within a tubular housing and having a thrust reverser having a stowed position fixed relative to the tubular housing and having a deployed position, and also having an existing first stage mixer which includes a number of first stage lobes enclosed in said tubular housing rearward from said jet engine, a method for cooling exhaust engine exhaust gases and for attenuating engine exhaust noise comprising the steps of: utilizing a second stage mixer coupled to the exterior of said tubular housing at an exhaust aperture of said tubular housing; the second stage mixer being unenclosed by a housing and having an exterior in contact with the ambient air to allow the air to enter the mixer and mix with the gas from the first stage mixer; the second stage mixer having a trailing edge forming the exit terminus for the gas flow; and said thrust reverser in the deployed position being positioned downstream of the exit terminus to reverse the gas flow; the second stage mixer receiving substantially all of the exhaust from the first stage mixer; utilizing a plurality of identically formed lobes and configuring said lobes (a) on their interior surfaces to force impinging hot gases from the jet engine towards the interior of the second stage mixer and (b) on their exterior surfaces to mix with ambient cooling air communicated from the exterior of said tubular housing to form gas cores which are smaller than those formed by the first stage mixer and to break the smaller gas cores into innumerable forms which are both cooler and noise attenuated.
20. The method according to claim 19 further comprising the steps of increasing the second stage mixer lobes in height from a circular configuration adjacent to the first stage mixer to an undulating configuration at the terminus of the second stage mixer.
21. In a nozzle for a jet engine enclosed within a tubular housing and having a thrust reverser having a stowed position fixed relative to the tubular housing and having a deployed position and also having an existing first stage mixer which includes a number of first stage lobes enclosed in said tubular housing rearward from said jet engine, a method for cooling hot exhaust engine exhaust gases and for attenuating engine exhaust noise comprising the steps of: utilizing a second stage mixer coupled to the exterior of said tubular housing at an exhaust aperture of said tubular housing; the second stage mixer being unenclosed by a housing and having an exterior in contact with the ambient air to allow the air to enter the mixer and mix with the gas from the first stage mixer; the second stage mixer having a trailing edge forming the exit terminus for the gas flow; and said thrust reverser in the deployed position being positioned downstream of the exit terminus to reverse the gas flow; the second stage mixer receiving substantially all of the exhaust from the first stage mixer; and utilizing a plurality of identically formed lobes.
22. The method according to claim 21 wherein the first stage lobes mix the hot exhaust gases with bypass air inside said tubular housing to provide a cooler and lower flow velocity mixture thereof, and in which said step of utilizing the plurality of identically formed lobes comprises the steps of further mixing the cooler and lower flow velocity mixture for further mixing thereof to level the disparate flow velocities attendant therewith, to reduce the peak velocities of the hot exhaust gases, and to increase flow of the cooler and lower flow velocity mixture whereby because noise is a function of jet exhaust velocity to the 7th power and because peak velocities from the hot exhaust gases are reduced, the engine exhaust noise is thereby reduced.
US11/027,982 2001-12-07 2005-01-04 Jet nozzle mixer Abandoned US20050247046A1 (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
US11/027,982 US20050247046A1 (en) 2001-12-07 2005-01-04 Jet nozzle mixer

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
US34036901P 2001-12-07 2001-12-07
US10/314,101 US6854260B2 (en) 2001-12-07 2002-12-07 Jet nozzle mixer
US11/027,982 US20050247046A1 (en) 2001-12-07 2005-01-04 Jet nozzle mixer

Related Parent Applications (1)

Application Number Title Priority Date Filing Date
US10/314,101 Continuation US6854260B2 (en) 2001-12-07 2002-12-07 Jet nozzle mixer

Publications (1)

Publication Number Publication Date
US20050247046A1 true US20050247046A1 (en) 2005-11-10

Family

ID=23333065

Family Applications (4)

Application Number Title Priority Date Filing Date
US10/314,101 Expired - Fee Related US6854260B2 (en) 2001-12-07 2002-12-07 Jet nozzle mixer
US11/027,981 Expired - Fee Related US7111448B2 (en) 2001-12-07 2005-01-04 Jet nozzle mixer
US11/027,982 Abandoned US20050247046A1 (en) 2001-12-07 2005-01-04 Jet nozzle mixer
US11/491,993 Expired - Fee Related US8069647B2 (en) 2001-12-07 2006-07-25 Jet nozzle mixer

Family Applications Before (2)

Application Number Title Priority Date Filing Date
US10/314,101 Expired - Fee Related US6854260B2 (en) 2001-12-07 2002-12-07 Jet nozzle mixer
US11/027,981 Expired - Fee Related US7111448B2 (en) 2001-12-07 2005-01-04 Jet nozzle mixer

Family Applications After (1)

Application Number Title Priority Date Filing Date
US11/491,993 Expired - Fee Related US8069647B2 (en) 2001-12-07 2006-07-25 Jet nozzle mixer

Country Status (7)

Country Link
US (4) US6854260B2 (en)
EP (1) EP1451461B8 (en)
AT (1) ATE358772T1 (en)
AU (1) AU2002351309A1 (en)
DE (1) DE60219345T2 (en)
HK (1) HK1066582A1 (en)
WO (1) WO2003050403A1 (en)

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20080242210A1 (en) * 2005-09-29 2008-10-02 Airbus Deutschland Gmbh Low-Noise Volume Flow Rate Throttling of Fluid-Carrying Pipes
US9863316B2 (en) 2013-08-14 2018-01-09 Ansaldo Energia Ip Uk Limited Intake arrangement in gas turbine power plant

Families Citing this family (62)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
ATE358772T1 (en) * 2001-12-07 2007-04-15 Jack H Anderson FLOWER MIXER FOR JET ENGINES
US7017331B2 (en) * 2002-12-07 2006-03-28 Anderson Jack H Jet nozzle mixer
US7482705B2 (en) * 2003-05-12 2009-01-27 Piercey Iii Gerald S Generator support plenum
FR2855558B1 (en) * 2003-05-28 2005-07-15 Snecma Moteurs TURBOMACHINE TUBE WITH NOISE REDUCTION
CA2547846A1 (en) * 2003-12-01 2005-11-17 The University Of Mississippi Method and device for reducing engine noise
US20080129053A1 (en) * 2004-05-12 2008-06-05 Piercey Gerald S Engine-generator set
US7418813B2 (en) * 2004-12-27 2008-09-02 General Electric Company Gas turbine engine exhaust nozzle including an infrared suppression system having a plurality of U-shaped blocking fins and method of assembling said exhaut nozzle
GB0505246D0 (en) * 2005-03-15 2005-04-20 Rolls Royce Plc Engine noise
US7216922B2 (en) * 2005-06-29 2007-05-15 The Boeing Company Fairing panel retainer apparatus
US20070000234A1 (en) * 2005-06-30 2007-01-04 Anderson Jack H Jet nozzle mixer
US7543452B2 (en) * 2005-08-10 2009-06-09 United Technologies Corporation Serrated nozzle trailing edge for exhaust noise suppression
DE602007004260D1 (en) * 2006-06-19 2010-03-04 Snecma Mixer for a twin-engine and corresponding nozzle and jet engine
FR2902758B1 (en) * 2006-06-21 2009-04-10 Airbus France Sas PROPELLANT AIRCRAFT ASSEMBLY COMPRISING AN EJECTION DUCT WITH A LEAK EDGE OFFSET
KR20080024623A (en) * 2006-09-14 2008-03-19 삼성전자주식회사 Display apparatus and data display method thereof
US8015794B2 (en) * 2006-10-20 2011-09-13 The Boeing Company Variable area flow duct employing secondary flows, and method therefor
US7966825B2 (en) * 2006-10-31 2011-06-28 Honeywell International Inc. Exhaust eductor system with a recirculation baffle
US20100028132A2 (en) * 2007-03-23 2010-02-04 Flodesign Wind Turbine Corporation Wind turbine with mixers and ejectors
US20110020107A1 (en) * 2007-03-23 2011-01-27 Flodesign Wind Turbine Corporation Molded wind turbine shroud segments and constructions for shrouds
US20110008164A1 (en) * 2007-03-23 2011-01-13 Flodesign Wind Turbine Corporation Wind turbine
US8657572B2 (en) 2007-03-23 2014-02-25 Flodesign Wind Turbine Corp. Nacelle configurations for a shrouded wind turbine
US8376686B2 (en) * 2007-03-23 2013-02-19 Flodesign Wind Turbine Corp. Water turbines with mixers and ejectors
US20090230691A1 (en) * 2007-03-23 2009-09-17 Presz Jr Walter M Wind turbine with mixers and ejectors
US20100316493A1 (en) * 2007-03-23 2010-12-16 Flodesign Wind Turbine Corporation Turbine with mixers and ejectors
US20110027067A1 (en) * 2007-03-23 2011-02-03 Flodesign Wind Turbine Corporation Coated shrouded wind turbine
US8021100B2 (en) * 2007-03-23 2011-09-20 Flodesign Wind Turbine Corporation Wind turbine with mixers and ejectors
US20110014038A1 (en) * 2007-03-23 2011-01-20 Flodesign Wind Turbine Corporation Wind turbine with skeleton-and-skin structure
US20100314885A1 (en) * 2007-03-23 2010-12-16 Flodesign Wind Turbine Corporation Shrouded wind turbine with rim generator and halbach array
US8622688B2 (en) * 2007-03-23 2014-01-07 Flodesign Wind Turbine Corp. Fluid turbine
US8726665B2 (en) 2007-06-05 2014-05-20 The Boeing Company Internal mixing of a portion of fan exhaust flow and full core exhaust flow in aircraft turbofan engines
US7762057B2 (en) * 2007-06-05 2010-07-27 The Boeing Company Internal mixing of a portion of fan exhaust flow and full core exhaust flow in aircraft turbofan engines
US20110215204A1 (en) * 2007-06-20 2011-09-08 General Electric Company System and method for generating thrust
US20080315042A1 (en) * 2007-06-20 2008-12-25 General Electric Company Thrust generator for a propulsion system
US20100019079A1 (en) * 2007-06-20 2010-01-28 General Electric Company Thrust generator for a rotary wing aircraft
WO2009110943A2 (en) * 2007-12-17 2009-09-11 Massachusetts Institute Of Technology Aircraft drag management structure
DE102007063018A1 (en) * 2007-12-21 2009-06-25 Rolls-Royce Deutschland Ltd & Co Kg Nozzle with guide elements
DE102008033984B4 (en) * 2008-07-21 2016-03-24 Friedrich Boysen Gmbh & Co. Kg exhaust system
EP2340199A4 (en) * 2008-09-29 2014-01-15 Andrew L Bender High efficiency turbine
EP2414673A4 (en) * 2009-03-30 2017-03-29 FloDesign Wind Turbine Corp. Segmented wind turbine
US8807458B2 (en) * 2009-04-29 2014-08-19 King Saud University Vortex-generating nozzle-end ring
CA2713317C (en) 2009-08-17 2017-09-26 Pratt & Whitney Canada Corp. Gas turbine engine exhaust mixer
DE102009047083C5 (en) * 2009-11-24 2013-09-12 J. Schmalz Gmbh Compressed air operated vacuum generator or vacuum gripper
US20110126512A1 (en) * 2009-11-30 2011-06-02 Honeywell International Inc. Turbofan gas turbine engine aerodynamic mixer
US8857761B2 (en) 2010-10-27 2014-10-14 Ata Engineering, Inc. Variable geometry aircraft pylon structure and related operation techniques
US8998169B2 (en) * 2011-02-28 2015-04-07 Control Components, Inc. Toothed gate valve seat
US10928101B2 (en) * 2011-06-10 2021-02-23 Carrier Corporation Ejector with motive flow swirl
US9567867B2 (en) 2011-09-14 2017-02-14 Ata Engineering, Inc. Methods and apparatus for deployable swirl vanes
CA2870290A1 (en) * 2012-04-10 2013-10-17 Rune Rubak Shrouded fluid turbine with boundary layer energising elements
JP5846617B2 (en) * 2012-07-24 2016-01-20 学校法人福岡大学 Fluid transfer device and fluid transfer method
EP2719620A3 (en) 2012-10-10 2017-10-18 Sikorsky Aircraft Corporation Upturned exhaust system for rotary wing aircraft
US9051047B2 (en) * 2012-11-12 2015-06-09 Sikorsky Aircraft Corporation Flexbeam rotor attachment to rotor blade
US10197010B2 (en) 2013-08-12 2019-02-05 The Boeing Company Long-duct, mixed-flow nozzle system for a turbofan engine
US9574518B2 (en) 2014-06-02 2017-02-21 The Boeing Company Turbofan engine with variable exhaust cooling
US10197009B2 (en) * 2014-07-30 2019-02-05 Pratt & Whitney Canada Corp. Gas turbine engine ejector
US10100705B2 (en) * 2014-08-27 2018-10-16 Sikorsky Aircraft Corporation Exhaust mixer and method of making same
US10001048B2 (en) 2016-03-21 2018-06-19 Paccar Inc Cyclonic thermal diffuser and method
CN106682287B (en) * 2016-12-08 2020-04-10 西北工业大学 Design method for molded lines of inner duct and outer duct of turbofan engine
US10927792B2 (en) * 2018-06-22 2021-02-23 The Boeing Company Jet noise suppressor
CN108894892A (en) * 2018-08-08 2018-11-27 西北工业大学 Petal-shaped noise reduction jet pipe
US11028778B2 (en) 2018-09-27 2021-06-08 Pratt & Whitney Canada Corp. Engine with start assist
US20210140369A1 (en) * 2019-11-13 2021-05-13 The Boeing Company Low pressure differential ejector pump utilizing a lobed, axisymmetric nozzle
GB2592267A (en) * 2020-02-24 2021-08-25 Altair Uk Ltd Pulse nozzle for filter cleaning systems
CN115030837B (en) * 2022-08-10 2022-11-08 中国空气动力研究与发展中心低速空气动力研究所 Spout noise reduction device

Citations (28)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3048975A (en) * 1959-11-19 1962-08-14 Rolls Royce Jet propulsion engines
US3084507A (en) * 1958-06-17 1963-04-09 Douglas Aircraft Co Inc Jet engine sound suppressor and reverser
US3568792A (en) * 1969-06-18 1971-03-09 Rohr Corp Sound-suppressing and thrust-reversing apparatus
US3572464A (en) * 1969-10-06 1971-03-30 Rohr Corp Method and apparatus for suppressing the noise of a fan-jet engine
US3601340A (en) * 1969-05-28 1971-08-24 Rohr Corp Method and apparatus for augmenting and reversing thrust and reducing nozzle base drag and noise of an aircraft jet engine
US3612212A (en) * 1969-08-11 1971-10-12 Rohr Corp Method and apparatus for suppressing the noise of a jet engine
US3696617A (en) * 1969-11-10 1972-10-10 Rohr Corp Turbo-fan propulsion apparatus and operating method
US3749316A (en) * 1972-06-26 1973-07-31 Rohr Industries Inc Sound suppressing thrust augmenting apparatus
US3776363A (en) * 1971-05-10 1973-12-04 A Kuethe Control of noise and instabilities in jet engines, compressors, turbines, heat exchangers and the like
US3982696A (en) * 1975-07-01 1976-09-28 Grumman American Aviation Corporation Jet noise suppressor nozzle
US4117671A (en) * 1976-12-30 1978-10-03 The Boeing Company Noise suppressing exhaust mixer assembly for ducted-fan, turbojet engine
US4254620A (en) * 1978-02-27 1981-03-10 The Boeing Company Jet engine multiduct noise suppressor
US4335801A (en) * 1980-12-15 1982-06-22 The Boeing Company Noise suppressing nozzle
US4422524A (en) * 1982-03-22 1983-12-27 Lockheed Corporation Variable shape, fluid flow nozzle for sound suppression
US4548034A (en) * 1983-05-05 1985-10-22 Rolls-Royce Limited Bypass gas turbine aeroengines and exhaust mixers therefor
US4592201A (en) * 1982-07-12 1986-06-03 General Electric Company Turbofan mixed flow exhaust system
US4666104A (en) * 1985-07-22 1987-05-19 Kelber Charles C Combination lift thrust device
US4813230A (en) * 1987-08-26 1989-03-21 Rolls-Royce Plc Gas turbine nozzles
US5127602A (en) * 1989-11-21 1992-07-07 Federal Express Corporation Noise reduction kit for jet turbine engines
US5157916A (en) * 1990-11-02 1992-10-27 United Technologies Corporation Apparatus and method for suppressing sound in a gas turbine engine powerplant
US5222359A (en) * 1991-06-28 1993-06-29 The Boeing Company Nozzle system and method for supersonic jet engine
US5440875A (en) * 1993-06-25 1995-08-15 United Technologies Corporation Fixed geometry mixer/ejector suppression system for turbofan aircraft engines
US5592813A (en) * 1995-07-06 1997-01-14 Avaero Hush kit for jet engine
US5706651A (en) * 1995-08-29 1998-01-13 Burbank Aeronautical Corporation Ii Turbofan engine with reduced noise
US5761900A (en) * 1995-10-11 1998-06-09 Stage Iii Technologies, L.C. Two-stage mixer ejector suppressor
US5884472A (en) * 1995-10-11 1999-03-23 Stage Iii Technologies, L.C. Alternating lobed mixer/ejector concept suppressor
US6233920B1 (en) * 1999-02-01 2001-05-22 Stage Iii Technologies, L.C. Contoured thrust reverser and lobed nozzle noise suppressor for gas turbine engines
US6854260B2 (en) * 2001-12-07 2005-02-15 Jack H. Anderson Jet nozzle mixer

Family Cites Families (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2123486B (en) * 1982-07-12 1986-01-02 Gen Electric Turbofan mixed flow exhaust system
GB2207468A (en) * 1987-06-01 1989-02-01 Secr Defence Vortex silencing in gas turbine engines
US5706900A (en) * 1996-11-19 1998-01-13 Liao; Dick Gardening tool for loosening soil

Patent Citations (29)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3084507A (en) * 1958-06-17 1963-04-09 Douglas Aircraft Co Inc Jet engine sound suppressor and reverser
US3048975A (en) * 1959-11-19 1962-08-14 Rolls Royce Jet propulsion engines
US3601340A (en) * 1969-05-28 1971-08-24 Rohr Corp Method and apparatus for augmenting and reversing thrust and reducing nozzle base drag and noise of an aircraft jet engine
US3568792A (en) * 1969-06-18 1971-03-09 Rohr Corp Sound-suppressing and thrust-reversing apparatus
US3612212A (en) * 1969-08-11 1971-10-12 Rohr Corp Method and apparatus for suppressing the noise of a jet engine
US3572464A (en) * 1969-10-06 1971-03-30 Rohr Corp Method and apparatus for suppressing the noise of a fan-jet engine
US3696617A (en) * 1969-11-10 1972-10-10 Rohr Corp Turbo-fan propulsion apparatus and operating method
US3776363A (en) * 1971-05-10 1973-12-04 A Kuethe Control of noise and instabilities in jet engines, compressors, turbines, heat exchangers and the like
US3749316A (en) * 1972-06-26 1973-07-31 Rohr Industries Inc Sound suppressing thrust augmenting apparatus
US3982696A (en) * 1975-07-01 1976-09-28 Grumman American Aviation Corporation Jet noise suppressor nozzle
US4117671A (en) * 1976-12-30 1978-10-03 The Boeing Company Noise suppressing exhaust mixer assembly for ducted-fan, turbojet engine
US4254620A (en) * 1978-02-27 1981-03-10 The Boeing Company Jet engine multiduct noise suppressor
US4335801A (en) * 1980-12-15 1982-06-22 The Boeing Company Noise suppressing nozzle
US4422524A (en) * 1982-03-22 1983-12-27 Lockheed Corporation Variable shape, fluid flow nozzle for sound suppression
US4592201A (en) * 1982-07-12 1986-06-03 General Electric Company Turbofan mixed flow exhaust system
US4548034A (en) * 1983-05-05 1985-10-22 Rolls-Royce Limited Bypass gas turbine aeroengines and exhaust mixers therefor
US4666104A (en) * 1985-07-22 1987-05-19 Kelber Charles C Combination lift thrust device
US4813230A (en) * 1987-08-26 1989-03-21 Rolls-Royce Plc Gas turbine nozzles
US5127602A (en) * 1989-11-21 1992-07-07 Federal Express Corporation Noise reduction kit for jet turbine engines
US5127602B1 (en) * 1989-11-21 1995-05-02 Federal Express Corp Noise reduction kit for jet turbine engines.
US5157916A (en) * 1990-11-02 1992-10-27 United Technologies Corporation Apparatus and method for suppressing sound in a gas turbine engine powerplant
US5222359A (en) * 1991-06-28 1993-06-29 The Boeing Company Nozzle system and method for supersonic jet engine
US5440875A (en) * 1993-06-25 1995-08-15 United Technologies Corporation Fixed geometry mixer/ejector suppression system for turbofan aircraft engines
US5592813A (en) * 1995-07-06 1997-01-14 Avaero Hush kit for jet engine
US5706651A (en) * 1995-08-29 1998-01-13 Burbank Aeronautical Corporation Ii Turbofan engine with reduced noise
US5761900A (en) * 1995-10-11 1998-06-09 Stage Iii Technologies, L.C. Two-stage mixer ejector suppressor
US5884472A (en) * 1995-10-11 1999-03-23 Stage Iii Technologies, L.C. Alternating lobed mixer/ejector concept suppressor
US6233920B1 (en) * 1999-02-01 2001-05-22 Stage Iii Technologies, L.C. Contoured thrust reverser and lobed nozzle noise suppressor for gas turbine engines
US6854260B2 (en) * 2001-12-07 2005-02-15 Jack H. Anderson Jet nozzle mixer

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20080242210A1 (en) * 2005-09-29 2008-10-02 Airbus Deutschland Gmbh Low-Noise Volume Flow Rate Throttling of Fluid-Carrying Pipes
US9863316B2 (en) 2013-08-14 2018-01-09 Ansaldo Energia Ip Uk Limited Intake arrangement in gas turbine power plant

Also Published As

Publication number Publication date
US20110265448A1 (en) 2011-11-03
EP1451461A1 (en) 2004-09-01
DE60219345D1 (en) 2007-05-16
US7111448B2 (en) 2006-09-26
WO2003050403A1 (en) 2003-06-19
US8069647B2 (en) 2011-12-06
ATE358772T1 (en) 2007-04-15
EP1451461B1 (en) 2007-04-04
US20050262826A1 (en) 2005-12-01
US20030196425A1 (en) 2003-10-23
EP1451461B8 (en) 2007-10-03
US6854260B2 (en) 2005-02-15
AU2002351309A1 (en) 2003-06-23
HK1066582A1 (en) 2005-03-24
DE60219345T2 (en) 2007-12-20

Similar Documents

Publication Publication Date Title
US6854260B2 (en) Jet nozzle mixer
US7251927B2 (en) Jet nozzle mixer
US20070000234A1 (en) Jet nozzle mixer
US7305817B2 (en) Sinuous chevron exhaust nozzle
US5638675A (en) Double lobed mixer with major and minor lobes
US5884472A (en) Alternating lobed mixer/ejector concept suppressor
US6292763B1 (en) Methods for designing variable cycle gas turbine engines
US7963099B2 (en) Fluted chevron exhaust nozzle
US5761900A (en) Two-stage mixer ejector suppressor
US4147027A (en) Thrust reverser nozzle
JP2000080958A (en) Nozzle with tab for jet noise reduction
US7810335B2 (en) Core exhaust mixer, having a variable area, for turbo-fan jet engines of supersonic aircraft
US20020164249A1 (en) Gas turbine engine exhaust nozzle
CA1099119A (en) Hybrid mixer for a high bypass ratio gas turbofan engine
US4227370A (en) By-pass gas turbine engines
Fishbach et al. NASA research in supersonic propulsion-A decade of progress
EP3597894A1 (en) Supersonic aircraft turbofan
Gutmark Duplex tab exhaust nozzle
Fogel et al. Avco Lycoming QCGAT program design cycle, demonstrated performance and emissions

Legal Events

Date Code Title Description
STCB Information on status: application discontinuation

Free format text: ABANDONED -- FAILURE TO RESPOND TO AN OFFICE ACTION