Tabbed nozzle for jet noise suppression

Abstract: The present invention relates to a gas turbine engine exhaust nozzle for suppressing jet noise. The nozzle 20 has an arrangement of tabs 44, 52 that extend in radially inward and outward directions. The tabs cause the formation of vortices which pull into mixing engagement the exhaust gas streams from the core and fan ducts and the ambient air. Tabs 48 that are continuous with the nozzle duct may be arranged between the angled tabs 44, 52. The tabs 44, 52 extend at an angle of up to 45°, more preferably up to 30°.

Description: This invention relates to gas turbine engine nozzles, and more particularly to nozzle constructions for jet noise suppression. Jet noise is created by the turbulent mixing of high velocity engine gases that emanate from the rear of a modern gas turbine. The turbulent mixing occurs between the high velocity gases and between the high velocity gases and ambient. The high velocity exhaust gases are typically a mixture of two sources - the hot gases resulting from the burnt fuel within the turbine's core flow (primary source) and cooler air discharged from fan bypass ducts (secondary source). The velocity of the core flow is typically in the order of 1600 ft/sec (488 ms -1 ) while the velocity of the fan bypass flow is typically in the order of 1000 ft/sec (305 ms -1 ). The two flows when fully mixed result typically in a mixed jet velocity of 1200 ft/sec (366 ms -1 ). The velocity gradient that exists at the different interfaces or shear regions, namely between the downstream mixed and the fan exhaust flows, between the fan exhaust flow and ambient, and between the core flow and ambient, results in flow disturbances. These flow disturbances or turbulence results in jet noise. The turbulent flow in the shear regions between the high velocity gases and the ambient air produce a significant component of the high levels of noise that are objectionable for aircraft operation from commercial airports. Due to the adverse impact noise has on the environment, many countries and airports have imposed increasingly strict noise reduction criteria on aircraft. In the United States, the Federal Aviation Administration (FAA) has imposed strict noise reduction limits on aircraft that are currently in use. In addition, the restrictions imposed by various airports range from financial penalties and schedule restrictions to an outright ban on the use of the aircraft. An effective and efficient noise reduction solution is necessary since these restrictions would severely cut short the useful life for certain types of aircraft that commercial airlines are currently using. Turbofan engines are categorized as either low bypass ratio or high bypass ratio, based on the ratio of bypass flow to core flow. Jet noise is a well-known problem with low bypass ratio engines. In the low bypass ratio jet engines, the exhaust gases emanating from the core and fan bypass ducts usually mix before they exit the engine's exhaust nozzle, where they form a high speed plume. The plume rips or shears against the slower ambient air as it rushes by creating flow turbulence and thus jet noise. Typically, newer jet engines are high bypass ratio engines which have lower (but still significant) levels of jet noise than low bypass ratio engines. Most high bypass ratio engines have separate flow nozzle exhaust systems. High bypass ratio engines have much larger fan flows, and overall larger total engine flow rates than the low bypass ratio engines. Thrust is obtained through larger mass flow rates, and lower jet velocities than low bypass ratio engines. Due to lower jet velocities, the level of jet noise is decreased in these high bypass ratio engines as compared to the low bypass ratio engines. However, jet noise remains a problem for modern high bypass ratio engines especially during operation at high engine power levels. High engine power is typically associated with aircraft take-off scenarios when the engine produces a high thrust and results in high velocity exhaust air. The FAA imposes strict noise requirements at high power. Modern, high bypass ratio engines have to comply with the requirement to provide ever-higher thrusts to power new and growth versions with higher takeoff gross weight of the aircraft. As a result, the modern, high bypass ratio engines operate at higher jet temperatures and pressure ratios and generate higher jet velocities and thus higher jet noise levels than earlier models of high bypass ratio engines. In the prior art of jet noise suppression, different structures have been devised to reduce noise. For example, a lobed mixer concept has been used in the past for the long duct, common flow exhaust systems such as those used in Pratt & Whitney's JT8D engine family. Examples of such noise suppression structures are found in U.S. Patent Nos. 4,401,269 and 5,638,675, both assigned to the Applicant. The '269 patent to Eiler and the '675 patent to Zysman et al disclose lobed mixers for a gas turbine engine. The lobed mixer includes axially and radially extending chutes. The chutes act as gas conduits whereby relatively cool, low velocity fan air is directed into the chutes and in turn into the hot, higher velocity core gas flow. The lobed mixer thus increases the mixing of the core and fan bypass gases. While the long duct, common flow exhaust systems of the prior art, as represented by the exhaust nozzles of the JT8D engine family, the '269 and '675 patents, have met with great commercial acceptance in the aerospace industry, the Applicant is constantly looking to improve the separate flow exhaust nozzle systems of gas turbine engines, especially during operation of the engines at high power levels. Other studies and nozzle configurations including tab concepts have been proposed and analyzed to understand the effects and physical phenomenon associated with the placement of tabs at the nozzle exit. However, heretofore nozzle configurations incorporating tabs for jet noise suppression have not resulted in a viable commercial product. Jet noise suppression improvements using nozzle tabs have to be lightweight, economical, easy to manufacture and incorporate in modern gas turbine engines. Further, nozzles incorporating tabs should not adversely impact engine thrust or performance. From a first aspect the invention provides a gas turbine exhaust nozzle as claimed in Claim 1. From a second aspect the invention provides a nozzle as claimed in claim 8. From a third aspect the invention provides an exhaust nozzle system as claimed in claim 9. In broad terms a preferred embodiment at the invention provides a gas turbine exhaust nozzle (20, 30) comprising an arrangement of tabs extending in radially inward and radially outward directions and extending circumferentially around the nozzle the tabs in use creating vortices for mixing the nozzle flow stream and an adjacent flow stream. According to a preferred embodiment of the present invention, there is provided a gas turbine engine exhaust nozzle for suppressing jet noise having an arrangement of nozzle tabs that are directed and extend in a radially inward and radially outward direction for increasing the effectiveness of the mixing process between exhaust gas streams and the ambient air. The nozzle tabs are disposed circumferentially on the exit of an exhaust nozzle. A preferred embodiment of the present invention comprises an alternating arrangement of tabs directed radially outwardly from the exhaust nozzle, tabs which are a smooth, continuous extension of the exhaust nozzle duct, tabs directed radially inwardly into the exhaust nozzle, followed by tabs which are a smooth, continuous extension of the exhaust nozzle duct. This alternating arrangement is repeated along the circumference of the nozzle exit. The present invention alters flow disturbances, which results in noise, by causing vortices to be set up by the tabs between the flow streams of the exhaust nozzles and ambient air. These vortices facilitate the mixing of the core and fan flow streams by drawing the fan flow stream radially inwardly into the core flow and alternatively by drawing the core flow radially outwardly into the fan flow, thus increasing the effectiveness of the mixing process. The nozzle tabs of the present invention are disposed such that there is a predetermined angular relationship between the tabs and the exhaust nozzle. The tab angles of protrusion radially into or out of the flow streams are determined to minimize the introduction of thrust losses for the level of noise reduction achieved. The height and number of tabs is a function of nozzle geometry. The present invention has utility in that it allows for jet noise suppression with minimal impact to engine thrust and performance. The present invention reduces jet noise close to the exit end of the engine due to the alteration of flow disturbances proximal to the exhaust nozzle. Further, the present invention has a minimal impact on the weight of the gas turbine engine. A preferred embodiment of the invention will now be described, by way of example only, with reference to the accompanying drawings, in which: Fig. 1 is a perspective view of the exit end of a gas turbine engine showing a view of a tabbed exhaust nozzle incorporating a preferred embodiment of the present invention; Fig. 2 is an end view showing a preferred tab arrangement of the present invention disposed on the primary exhaust nozzle of Fig. 1; Fig. 2A is a sectional view taken along the line 2A-2A of Fig. 2; Fig. 2B is a sectional view taken along the line 2B-2B of Fig. 2; Fig. 2C is a sectional view taken along the line 2C-2C of Fig. 2; Fig. 3 is a view similar to Fig. 2 showing the interaction between the gas streams that exit the exhaust nozzle and the ambient air; Fig. 4 is schematic view, partly broken away and partly in section, of the exit end of the exhaust nozzle showing further details of the interaction between the gas flow streams caused by the preferred tab arrangement of the present invention. Referring to Fig. 1, in a typical turbofan engine 10, having a longitudinally extending central axis, air enters the upstream end of the engine. As is well known in the art and, therefore, not illustrated herein, as air enters the front of the jet engine it passes through the fan and is split between a primary or core flow and a secondary or bypass flow. The primary flow first enters a low pressure compressor then a high pressure compressor. The air is then mixed with fuel in a combustion chamber and the mixture is ignited and burned, thereby increasing its pressure and temperature. The resultant combustion products then flow into a high pressure turbine and a low pressure turbine which extract energy from the combustion gases to turn the fan and compressor. The gases then expand through an inner exhaust nozzle 20 to produce useful thrust. The bypass stream is compressed by the fan, flows outside the core of the engine through an annular duct concentric with the core engine and is exhausted through an outer exhaust nozzle 30 as additional useful thrust. The two concentric flow streams join downstream of the turbine exhaust area of the engine. The two flows then mix together and with the surrounding ambient flow as described hereinafter. Referring to Fig. 2, the preferred embodiment of the present invention is an arrangement of alternating tabs 40. This arrangement of tabs includes tabs 44 that are directed radially outward and extend into the fan flow stream, tabs 48 which are a smooth, continuous extension of the exhaust nozzle duct 20, tabs 52 that are directed radially inward and extend into the core flow stream, followed by tabs 48 which are a smooth continuous extension of the exhaust nozzle duct. This alternating arrangement of tabs is repeated in sequence and is disposed along the entire circumference of the nozzle 20 exit. Referring to Fig. 2A and 2C, the tabs 44, 52 directed radially outwardly and inwardly with respect to the core and fan flow streams have an angular relationship with respect to the gas flow streams exiting through the nozzles 20, 30. The tabs protrude at predetermined angles into either the core or fan flow stream. The angular orientation for the nozzle tabs of the preferred embodiment were determined to minimize any adverse impact to engine thrust. In the exemplary embodiment of the present invention, the angle of protrusion can range from zero to forty-five degrees (45°). Angles beyond 45° adversely impact engine thrust. Referring to Fig. 3, the vortices 56 created by the tabs 40 on the exhaust nozzle pull the core flow stream and fan flow stream into mixing engagement. The tabs 44 that are directed radially outwardly and extend into the fan flow stream create vortices that cause the core flow emanating from inner exhaust nozzle 20 to expand or be drawn into the secondary flow emanating from outer exhaust nozzle 30. The tabs 52 that are directed radially inwardly and extend into the core flow create vortices that cause the fan flow to penetrate or be drawn into the core flow. The tabs 48 that are continuous extensions of the exhaust nozzle prevent the outwardly and inwardly directed core and fan flows from interfering with each other. Referring to Figs. 3 and 4, each tab 44 that is directed radially outwardly produces a pair of streamwise vortices that rotate in opposite directions. The vortices cause the mixing of the core flow and fan flow. Effective mixing between the two flow streams alters flow disturbances between the core flow and fan flow interfaces and results in attenuating jet noise. The right edge 70 of tab 44 creates a vortex having a flow direction which is counter-clockwise. This vortex pulls the core flow outwardly into the fan flow. The left edge 78 of the tab 44 creates a vortex having a flow direction which is clockwise. This latter vortex pulls the core flow outwardly into the fan flow. Similarly, a pair of streamwise vortices that rotate in opposite directions are created by each tab 52 that is directed radially inwardly. The vortices created by the tabs 52 draw the fan flow radially inwardly into the core flow, thus pulling the fan flow and core flow into mixing engagement. The strength of this pair of vortices is approximately equal to the circulation or degree of mixing defined according to the following equation: circulation = 2 U h tan∝ where "U" is the velocity difference between the velocity of the core flow and the fan flow, " h " is the height of a tab (the distance from the base of the tab to its downstream end) and alpha (α) is the angle of protrusion radially inwardly or outwardly of the flow stream. Thus, the circulation of the exhaust flows is a function of the difference of the velocities of flow streams that are mixing, the height or protrusion of the tabs into the flow streams and the tangent of the angle of protrusion of the tabs into or out of the flow streams. The angle of protrusion into the flow streams may be in a range of zero degrees to forty-five degrees (45°). Angles greater than forty-five degrees may add appreciably to the thrust losses due to local flow separation which is detrimental to the performance of the nozzle. Even though the angular orientation of the tabs can range from 0 to 45°, for practical purposes the angular orientation range is from 0 to 30°. In accordance with the triangular embodiment of the present invention, the height of the tabs is a function of the circumferential size of the tabs. The maximum height of the triangular tabs can be represented by the following equation: h = ((Πd)/N)/2 where Πd is the circumferential size of the tabs (the circumference of the nozzle at the base of the tabs) and N is the number of tabs disposed around the exit nozzle. In the turbofan engine industry, the effect of noise on humans is expressed in terms of an effective perceived noise level, EPNdB, expressed in decibels. At high mixed jet velocities in the range of 1200 feet per second (366 ms -1 ), the jet noise suppression achieved by the present invention is over three EPNdB (3 EPNdB). These higher jet velocities are achieved at take-off when noise suppression is desired. Thus, the present invention provides the required level of jet noise suppression during engine operation at high power levels. The tabbed nozzle of the present invention may be used in engines having both a primary core and a secondary fan nozzle as well as in engines having only a primary core exhaust nozzle. For bypass engines, the tab arrangement of the present invention can be disposed either on the exit end of both the primary and secondary nozzle or on any one of the two nozzles. The tabbed nozzle of the present invention may be implemented in a variety of ways by any of a number of arrangements of tabs extending in radially inward and outward directions with a substantial angular offset from one another. As described hereinabove, the tabs may form an alternating arrangement as disclosed or alternatively some other combination of radially directed tabs and tabs continuous with the surface of the nozzle. Other combinations of the tabs may include a plurality of radially inwardly directed tabs adjacent to a plurality of radially outwardly directed tabs. The tabs may be spaced from each other or may be adjacent each other. It will also be understood by those skilled in the art that the above number of tabs described are derived for particular exhaust nozzle size and geometries. The number of tabs may be adjusted to suit different nozzles. From the above it will be seen there has been described an arrangement leading to jet noise suppression, especially during engine operation at high power levels, without the addition of appreciable thrust losses, which results in minimum weight penalty for the gas turbine engine. All of the foregoing embodiments are representative of the preferred embodiment. It suffices for the present invention that a gas turbine engine exhaust nozzle for suppressing jet noise includes an arrangement of nozzle tabs that are directed and extend in radially inward and outward directions, for mixing the exhaust flow streams and the ambient. Although the invention has been shown and described with respect to detailed embodiments thereof, it should be understood by those skilled in the art that various changes in form and detail thereof may be made without departing from the scope of the claimed invention.

Claims: A gas turbine exhaust nozzle (20, 30) comprising an arrangement of tabs (44, 52) extending in radially inward and radially outward directions and extending circumferentially around the nozzle, characterised in that said tabs are triangular in shape and that said tabs in use create vortices (56) for mixing the nozzle flow stream and an adjacent flow stream. The nozzle of Claim 1, wherein the arrangement of tabs comprises a repeated sequence of tabs, wherein the sequence includes tabs (44) directed radially outward from the exhaust nozzle, tabs (52) directed radially inward from the exhaust nozzle and continuous tabbed extensions (48) of the nozzle therebetween. The nozzle of Claim 1 or 2, wherein the tabs (44, 52) directed radially inward and outward have a substantial angular offset from one another. The nozzle of Claim 3, wherein the tabs (44, 52) directed radially inward and outward protrude into or out of the nozzle flow stream at an angle of up to forty-five degrees (45°). The nozzle of Claim 4, wherein the tabs (44, 52) directed radially inward and outward protrude into or out of the nozzle flow stream at an angle of up to thirty degrees (30°). The nozzle of Claim 5 wherein the arrangement of triangular-shaped tabs (44, 52) having a circumferential size, further comprises tab orientations to maximize the degree of mixing of flow streams, the degree of mixing being a function of the product of the velocity difference between the inner and outer flow streams, the height of the tabs and tangent of the angle of protrusion of the tabs into or out of the flow stream. The nozzle of Claim 6, wherein the height of the tabs further comprises a maximum size that is limited by the ratio of the circumferential size of the tabs and the number of tabs disposed along the nozzle. A nozzle (20) having an exit end in a gas turbine engine for suppressing jet noise comprising: a tab (44) directed radially outward and disposed on the exit end of the nozzle; a tab (52) directed radially inward and disposed on the exit end of the nozzle; and a tab (48), continuous with the exit end of the nozzle therebetween; wherein the alternating arrangement of tabs is disposed circumferentially about the exit end of the nozzle, characterised in that said tab directed radially outward and said tab directed radially inward are triangular in shape; and wherein the tabs create vortices which pull into mixing engagement the nozzle flow stream and a portion of ambient air. An exhaust nozzle system in a gas turbine engine for suppressing jet noise, the engine having a longitudinally extending central axis, concentric inner and outer flow paths for carrying inner and outer fluid flow streams through the engine, an inner and outer nozzle to discharge the inner and outer flow streams to ambient, characterised in that at least one of the nozzles comprises: an arrangement of tabs (44,52) directed and extending in radially inward and outward directions, the tabs disposed circumferentially about the nozzle, wherein the tabs are triangular in shape and create vortices (56) which pull into mixing engagement the nozzle flow streams and a portion of ambient air. A gas turbine engine comprising a nozzle as claimed in any preceding claim. An engine as claimed in claim 10 wherein said engine comprises an inner nozzle (20) and an outer nozzle (30) wherein one or both of said nozzles is a nozzle as claimed in any of claims 1 to 10.

Inventors: LOW JOHN K C, MATHEWS DOUGLAS C, LOW, JOHN K.C., MATHEWS, DOUGLAS C.
Applicants: UNITED TECHNOLOGIES CORP, UNITED TECHNOLOGIES CORPORATION
Application year: 1999
Main classification: F02K1
IPC codes: F02K1