US2657575A - Asymmetric adjustable supersonic nozzle - Google Patents

Asymmetric adjustable supersonic nozzle Download PDF

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US2657575A
US2657575A US769968A US76996847A US2657575A US 2657575 A US2657575 A US 2657575A US 769968 A US769968 A US 769968A US 76996847 A US76996847 A US 76996847A US 2657575 A US2657575 A US 2657575A
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nozzle
test section
section
wall
walls
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Harry J Allen
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K1/00—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
    • F02K1/06—Varying effective area of jet pipe or nozzle
    • F02K1/09—Varying effective area of jet pipe or nozzle by axially moving an external member, e.g. a shroud
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles

Definitions

  • This invention relates to improvements in nozzles, and more particularly to improvements in supersonic nozzles.
  • the principal object of this invention is to provide a nozzle in which the flow Mach number is continuously adjustable.
  • Another object of this invention is to provide a nozzle in which the flow passage is asymmetncal and adjustable to define various Mach numbers.
  • Fig. 1 is an isometric drawing of a nozzle embodying my invention and shows a wall broken away at the inlet end of the nozzle to better illustrate the construction.
  • Fig. 2 is a longitudinal section of Fig. 1 showing the nozzle conventionally positioned relative to a source of fluid flow l1 and supported as at [8 to permit longitudinal movement of wall [5.
  • the area of the passage upstream of the test section In a wind tunnel in order to provide a supersonic Mach number at the section of the flow passage in which the model undergoing tests is located, known as the test section, the area of the passage upstream of the test section must be varied in an amount dependent upon the Mach number desired at th test section.
  • the ratio of the area of the test section to the area of the throat, or minimum section, upstream of the test section uniquely determines the Mach number at the test section.
  • the other method has been to provide the nozzle section upstream of the test section with one or more walls which could be flexed by using a series of jacks to vary the ratio of test section to throat section area and so vary the test section Mach number.
  • This method sufiers from mechanical complexity resulting from the required system of jacks necessary for the manipulation of the flexible section of the nozzle.
  • the passage is asymmetric andthe construction of the nozzle is such'that the required adjustment of area ratio necessary to change the Mach number can be accomplished by sliding one of the walls of the nozzle longitudinally, or nearly longitudinally, relative to the other walls.
  • This type of nozzle affords a continuously adjustabl Mach number at the test section, an advantage not obtained by the fixed nozzle meth od previously discussed, and possesses greater mechanical and structural simplification as compared with the flexible-wall type of nozzle.
  • My nozzle can find application in steam or gas turbines, in the exit of jet motors and in supersonic wind tunnel test sections.
  • the reference numeral I0 designates a tubular body portion having a, restricted, asymmetric flow passage ll defined by parallel side plates 12 and I3, inwardly curved top plate Hi, and outwardly flared bottom plate I5.
  • Bottom flared plate I5 is in sliding engagement with parallel sides I 2 and I3, and is movable longi-- 3 tudinally in guides l6 which are secured to the side plates by spot welding or other suitable means.
  • the fluid from the source enters the inlet or upstream end of the nozzle in the directionindicatediby the arrow H-
  • the test section i. e., the section wherein a model is mounted for studying aerodynamic flow
  • the test section is located at the opposite or downstream end: of the nozzle.
  • the adjustable wall or plate With the adjustable wall or plate at the position indicated by the dotted line I5, the area ratio is high and accordingly the Mach number at the test section, or downstream end; isv low, Progressive forward motion of the movable plate to a position shown by the solid line l5, decreases the area ratio and, therefore, increases the Mach. number at the test section. It is apparent that either the plate 14 or the plate l may be fixed, the opposite plate being, adjustable.
  • asymmetrical supersonic nozzle for use with a wind tunnel saiId' nozzle having. an inlet end and a, testsection. at the opposite end and comprising. a. pair of'fixed. parallel walls, a fixed curved wall, andlamovable wall having a. curvature substantially inverse to that of said curved wall. so. as to form aflared. inlet end for said nozzle, said. walls defining a passage of substan-- tially. rectangular cross-section, the movement of said. movable wall permitting: theadjustment of the. relative areas within. said passage to obtain a continuous adjustment of the Mach number at the test section of said'passage-the upper limit of the ratio of the cross sectional areas: at the inlet end. and test section being unity.
  • Anasymmetrical supersonic nozzle for use with a: wind. tunnel saidnozzle having an: inlet endandea test section at the opposite. end, and. comprising a pair offixed parallelwalls, afixedwall contiguous withsaid parallel walls-andhaV- ing, a curved, inlet endportion, a longitudinally movable wall contiguous with said parallel; walls and positioned opposite said curved wall; saidmovable wall havingan inletend-portion. of re.- verse. curvature withrespect to the first-men:- tionedcurved portion whereby longitudinal movement of said movablewall varies the relative-areas within the passagedefinedby said walls to there-- by adjust the Mach number atthe test sectionE of said nozzle, theupper. limit of the, ratioof th cross-sectional areas at. the inlet endandthe. test sectionbeing unity.
  • An asymmetrical supersonic nozzle suitable for use with a wind tunnel, said nozzle having an inlet end and a test section at the opposite end and comprising a pair of fixed parallel walls, a fixed wall having a concave inner surface at the inlet end of the. nozzle, a longitudinally translatable wall having; a convex inner surface at the inlet" end, said four walls forming a passage, the cross-sections of which substantially define rectangles, guide means for the longitudinally translatable wall; said' concave and convex surfaces of the walls being relatively movable to vary the ratio of the area of the passage in the inlet portion with respect to the area of the test section, andhencethe Mach number at the test section.
  • asymmetrical supersonic nozzle having an inlet end and a test section at the opposite end thereof and comprising four walls the inner'surfaces of which form substantially rectangular sections along the length of the nozzle, two oppositewalls being fixed, and the other'two opposite walls, being relatively movable longitudinally of each other, a guide means. for a movable wall, said relatively movable walls each having an" oppositely curved portion. at the inlet end of the nozzle for varying. the cross-section area of the passageway near the inlet end, and said cross section area of the inlet end n'ever' exceeding the cross-sectionarea of the test section of the nozzle inany adjustedlpositionof the nozzle.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Aerodynamic Tests, Hydrodynamic Tests, Wind Tunnels, And Water Tanks (AREA)

Description

Nov. 3, 1953 H. J. ALLEN ASYMMETRIC ADJUSTABLE SUPERSONIC NOZZLE Fil ed Aug. 21, 1947 FIG.
FIG. 2
:lwuC/H fo-n HARRY J. ALLEN Patented Nov. 3, 1953 OFFICE ASYIVIMETRIC ADJUSTABLE SUPERSONIC NOZZLE Harry J. Allen, Palo Alto, Calif.
Application August 21, 1947, Serial No. 769,968
(Cl. 73147) I (Granted under Title 35, U. S. Code (1952),
Claims.
sec. 266) This invention relates to improvements in nozzles, and more particularly to improvements in supersonic nozzles.
The principal object of this invention is to provide a nozzle in which the flow Mach number is continuously adjustable.
Another object of this invention is to provide a nozzle in which the flow passage is asymmetncal and adjustable to define various Mach numbers.
In the accompanying drawings, forming a part of this specification, and in which like numerals are employed to designate like parts throughout the same,
Fig. 1 is an isometric drawing of a nozzle embodying my invention and shows a wall broken away at the inlet end of the nozzle to better illustrate the construction.
Fig. 2 is a longitudinal section of Fig. 1 showing the nozzle conventionally positioned relative to a source of fluid flow l1 and supported as at [8 to permit longitudinal movement of wall [5.
No claim is made to a specific source of fluid flow or to a particular supporting structure, these being of any convenient known construction,
In a wind tunnel in order to provide a supersonic Mach number at the section of the flow passage in which the model undergoing tests is located, known as the test section, the area of the passage upstream of the test section must be varied in an amount dependent upon the Mach number desired at th test section. The ratio of the area of the test section to the area of the throat, or minimum section, upstream of the test section uniquely determines the Mach number at the test section.
There have been, in general, two methods employed in the past for changing the Mach number in supersonic wind tunnels. One method has been to employ interchangeable nozzles designed to attain specific Mach numbers at the test section. This method suffers from the disadvantages that as many nozzles as desired test section Mach numbers must be constructed, and that the Mach number cannot be easily varied through small increments.
The other method has been to provide the nozzle section upstream of the test section with one or more walls which could be flexed by using a series of jacks to vary the ratio of test section to throat section area and so vary the test section Mach number. This method sufiers from mechanical complexity resulting from the required system of jacks necessary for the manipulation of the flexible section of the nozzle.
In addition, overstressing of the flexible section can easily occur by improper jack operation thus resulting in permanent set or structural failure. In order to prevent such an occurrence it is necessary to employ elaborate safety devices thereby increasing the cost of production and maintenance.
In the new type of nozzl of the present invention, the passage is asymmetric andthe construction of the nozzle is such'that the required adjustment of area ratio necessary to change the Mach number can be accomplished by sliding one of the walls of the nozzle longitudinally, or nearly longitudinally, relative to the other walls. This type of nozzle affords a continuously adjustabl Mach number at the test section, an advantage not obtained by the fixed nozzle meth od previously discussed, and possesses greater mechanical and structural simplification as compared with the flexible-wall type of nozzle.
Large scale tests of asymmetric nozzles indicate that, with suitable design, satisfactory aerodynamic characteristics, such as shock wave, compression ratio, boundary layer growth, etc., can be obtained which give satisfactory flow throughout the entire desired range of Mach number setting for all positions of the adjustable wall. If the flow at all stations is less than unity i. e. subsonic, it is well known that the Mach number increases as the area decreases. The reverse is however true at supersonic speeds and it can be shown mathematically that for a supersonic sound tunnel the speed of the test section is uniquely determined by the ratio of the testsection area to the minimum or throat area ahead of it. Furthermore, this adjustable wall may be moved parallel to the high speed section walls, thus maintaining a constant section height at the high speed section. The direction of motion can, however, be varied somewhat so that a greater or lesser height of passage is obtained at th high speed section as the Mach number is changed.
My nozzle can find application in steam or gas turbines, in the exit of jet motors and in supersonic wind tunnel test sections.
In the drawings, wherein for the purpose of illustration, is shown a preferred embodiment of my invention, the reference numeral I0 designates a tubular body portion having a, restricted, asymmetric flow passage ll defined by parallel side plates 12 and I3, inwardly curved top plate Hi, and outwardly flared bottom plate I5. Bottom flared plate I5 is in sliding engagement with parallel sides I 2 and I3, and is movable longi-- 3 tudinally in guides l6 which are secured to the side plates by spot welding or other suitable means.
In operation, as indicated in Fig. 2, the fluid from the source, schematically represented at ll, enters the inlet or upstream end of the nozzle in the directionindicatediby the arrow H- When the nozzle is employed in a wind tunnel, the test section, i. e., the section wherein a model is mounted for studying aerodynamic flow, is located at the opposite or downstream end: of the nozzle. With the adjustable wall or plate at the position indicated by the dotted line I5, the area ratio is high and accordingly the Mach number at the test section, or downstream end; isv low, Progressive forward motion of the movable plate to a position shown by the solid line l5, decreases the area ratio and, therefore, increases the Mach. number at the test section. It is apparent that either the plate 14 or the plate l may be fixed, the opposite plate being, adjustable.
The inventiondescribed herein. may be manufactured and" used by or for the Government of the United States of America for governmental purposes without the payment of. any royalties thereon, or therefor.
What is claimed is:
1.. asymmetrical supersonic nozzle for use with a wind tunnel saiId' nozzle having. an inlet end and a, testsection. at the opposite end and comprising. a. pair of'fixed. parallel walls, a fixed curved wall, andlamovable wall having a. curvature substantially inverse to that of said curved wall. so. as to form aflared. inlet end for said nozzle, said. walls defining a passage of substan-- tially. rectangular cross-section, the movement of said. movable wall permitting: theadjustment of the. relative areas within. said passage to obtain a continuous adjustment of the Mach number at the test section of said'passage-the upper limit of the ratio of the cross sectional areas: at the inlet end. and test section being unity.
2. Anasymmetrical supersonic nozzle for use with a: wind. tunnel saidnozzle having an: inlet endandea test section at the opposite. end, and. comprising a pair offixed parallelwalls, afixedwall contiguous withsaid parallel walls-andhaV- ing, a curved, inlet endportion, a longitudinally movable wall contiguous with said parallel; walls and positioned opposite said curved wall; saidmovable wall havingan inletend-portion. of re.- verse. curvature withrespect to the first-men:- tionedcurved portion whereby longitudinal movement of said movablewall varies the relative-areas within the passagedefinedby said walls to there-- by adjust the Mach number atthe test sectionE of said nozzle, theupper. limit of the, ratioof th cross-sectional areas at. the inlet endandthe. test sectionbeing unity.
3. An asymmetrical supersonic nozzle suitable for use with a wind tunnel, said nozzle having an inlet end and a test section at the opposite end and comprising a pair of fixed parallel walls, a fixed wall having a concave inner surface at the inlet end of the. nozzle, a longitudinally translatable wall having; a convex inner surface at the inlet" end, said four walls forming a passage, the cross-sections of which substantially define rectangles, guide means for the longitudinally translatable wall; said' concave and convex surfaces of the walls being relatively movable to vary the ratio of the area of the passage in the inlet portion with respect to the area of the test section, andhencethe Mach number at the test section.
4. The combination set forth in claim 3 further definedin that a continuous adjustment is provided. by movement of the longitudinally transla-table wall and that the ratio of the minimum cross-section area of the inlet end and test section does not exceed unity.
5. asymmetrical supersonic nozzle having an inlet end and a test section at the opposite end thereof and comprising four walls the inner'surfaces of which form substantially rectangular sections along the length of the nozzle, two oppositewalls being fixed, and the other'two opposite walls, being relatively movable longitudinally of each other, a guide means. for a movable wall, said relatively movable walls each having an" oppositely curved portion. at the inlet end of the nozzle for varying. the cross-section area of the passageway near the inlet end, and said cross section area of the inlet end n'ever' exceeding the cross-sectionarea of the test section of the nozzle inany adjustedlpositionof the nozzle.
HARRY J ALLEN.
References Cited in the file of this patent- UNITED STATES PATENTS Number Name Date 1,376,178 Wage'nse'il Apr. 26, 1921 2,171,817" Wayner et a1. Sept. 5, 1939 2,342,262 Franz et--a1i-.. 'g Feb; 22,1944 2,405g'723 Way Aug. 13, 1946' 2,41'85488 Th'omp'son Apr. 8; 1947 2540594 Price- Feb. 6, 1951' 2,5701629 Anxion'naze't all Oct. 9, 1951 2 59334 20 Diehl Apr. 22,1952. 2,608,053 Davidson Aug. 26, 1952 FOREIGN PATENTS Number Country Date 58.0;995: Great Britain Sept. 26, 1946 OTHER REFERENCES Ser. No. 326,141, Ramshorn-(A'. P.-C.) published May. 11,. I943.
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Cited By (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2812980A (en) * 1950-06-16 1957-11-12 Snecma Jet deflecting device
DE1061140B (en) * 1956-02-17 1959-07-09 Soc D Forges Et Ateliers Du Cr Rectangular Laval control nozzle for a flow channel, especially a high-speed wind channel
US2938334A (en) * 1957-12-23 1960-05-31 United Aircraft Corp Supersonic inlet
US2947354A (en) * 1956-02-17 1960-08-02 Creusot Forges Ateliers Movable panel
US3019601A (en) * 1959-10-05 1962-02-06 United Aircraft Corp Pen-shaped nozzle with thrust deflector
US3020714A (en) * 1956-07-03 1962-02-13 Snecma Device for controlling the jet of a reaction propulsion motor
US3038305A (en) * 1953-01-23 1962-06-12 Lockheed Aircraft Corp Subsonic, supersonic propulsive nozzle
US3111842A (en) * 1960-07-19 1963-11-26 Cook Electric Co Supersonic wind tunnel
DE1175036B (en) * 1960-07-01 1964-07-30 Rolls Royce Aircraft jet engine plant
US20040226379A1 (en) * 2003-02-20 2004-11-18 Michael Ochwat Method and device for checking the air noise of a motor vehicle
US20090206207A1 (en) * 2005-04-18 2009-08-20 Scott Rethorst Supersonic aircraft footprint spreading control system and method
CN105464838A (en) * 2014-09-25 2016-04-06 波音公司 Methods and apparatus for passive thrust vectoring and plume deflection
CN110450964A (en) * 2018-05-07 2019-11-15 南京普国科技有限公司 Class axial symmetry tilt outlet, which is received, expands jet pipe and its design method

Citations (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US1376178A (en) * 1916-01-24 1921-04-26 Hugo Junkers Radiator
US2171817A (en) * 1936-12-22 1939-09-05 Messerschmitt Boelkow Blohm Radiator for aviation engines
US2342262A (en) * 1939-05-30 1944-02-22 Franz Anselm Adjustable reaction nozzle
US2405723A (en) * 1946-08-13 Propulsion apparatus
GB580995A (en) * 1944-01-21 1946-09-26 British Thomson Houston Co Ltd Improvements in and relating to aircraft employing jet propulsion
US2418488A (en) * 1944-07-29 1947-04-08 Westinghouse Electric Corp Power-plant apparatus
US2540594A (en) * 1946-08-23 1951-02-06 Lockheed Aircraft Corp Ram jet engine having variable area inlets
US2570629A (en) * 1945-10-05 1951-10-09 Anxionnaz Adjustable pipe for the intake of air and expansion of the driving gases in reactionjet propellers for projectiles and vehicles
US2593420A (en) * 1946-05-28 1952-04-22 Walter S Diehl Variable area nozzle
US2608053A (en) * 1946-05-03 1952-08-26 Davidson Milton Variable area diffuser or effuser

Patent Citations (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2405723A (en) * 1946-08-13 Propulsion apparatus
US1376178A (en) * 1916-01-24 1921-04-26 Hugo Junkers Radiator
US2171817A (en) * 1936-12-22 1939-09-05 Messerschmitt Boelkow Blohm Radiator for aviation engines
US2342262A (en) * 1939-05-30 1944-02-22 Franz Anselm Adjustable reaction nozzle
GB580995A (en) * 1944-01-21 1946-09-26 British Thomson Houston Co Ltd Improvements in and relating to aircraft employing jet propulsion
US2418488A (en) * 1944-07-29 1947-04-08 Westinghouse Electric Corp Power-plant apparatus
US2570629A (en) * 1945-10-05 1951-10-09 Anxionnaz Adjustable pipe for the intake of air and expansion of the driving gases in reactionjet propellers for projectiles and vehicles
US2608053A (en) * 1946-05-03 1952-08-26 Davidson Milton Variable area diffuser or effuser
US2593420A (en) * 1946-05-28 1952-04-22 Walter S Diehl Variable area nozzle
US2540594A (en) * 1946-08-23 1951-02-06 Lockheed Aircraft Corp Ram jet engine having variable area inlets

Cited By (18)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2812980A (en) * 1950-06-16 1957-11-12 Snecma Jet deflecting device
US3038305A (en) * 1953-01-23 1962-06-12 Lockheed Aircraft Corp Subsonic, supersonic propulsive nozzle
DE1061140B (en) * 1956-02-17 1959-07-09 Soc D Forges Et Ateliers Du Cr Rectangular Laval control nozzle for a flow channel, especially a high-speed wind channel
US2947354A (en) * 1956-02-17 1960-08-02 Creusot Forges Ateliers Movable panel
US3020714A (en) * 1956-07-03 1962-02-13 Snecma Device for controlling the jet of a reaction propulsion motor
US2938334A (en) * 1957-12-23 1960-05-31 United Aircraft Corp Supersonic inlet
US3019601A (en) * 1959-10-05 1962-02-06 United Aircraft Corp Pen-shaped nozzle with thrust deflector
DE1175036B (en) * 1960-07-01 1964-07-30 Rolls Royce Aircraft jet engine plant
DE1175036C2 (en) * 1960-07-01 1965-02-04 Rolls Royce Aircraft jet engine plant
US3111842A (en) * 1960-07-19 1963-11-26 Cook Electric Co Supersonic wind tunnel
US20040226379A1 (en) * 2003-02-20 2004-11-18 Michael Ochwat Method and device for checking the air noise of a motor vehicle
US7036361B2 (en) * 2003-02-20 2006-05-02 Dr. Ing H.C.F. Porsche Aktiengesellschaft Method and device for checking the air noise of a motor vehicle
US20090206207A1 (en) * 2005-04-18 2009-08-20 Scott Rethorst Supersonic aircraft footprint spreading control system and method
US7861966B2 (en) * 2005-04-18 2011-01-04 Vehicle Research Corporation Supersonic aircraft footprint spreading control system and method
CN105464838A (en) * 2014-09-25 2016-04-06 波音公司 Methods and apparatus for passive thrust vectoring and plume deflection
CN105464838B (en) * 2014-09-25 2019-05-21 波音公司 Method and apparatus for being deflected by dynamicthrust guiding and plume
CN110450964A (en) * 2018-05-07 2019-11-15 南京普国科技有限公司 Class axial symmetry tilt outlet, which is received, expands jet pipe and its design method
CN110450964B (en) * 2018-05-07 2020-11-24 南京普国科技有限公司 Axisymmetric inclined outlet convergent-divergent nozzle and design method thereof

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