WO1996012881A1 - Buse d'evacuation d'air de ventilation, a section de superficie variable - Google Patents

Buse d'evacuation d'air de ventilation, a section de superficie variable Download PDF

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Publication number
WO1996012881A1
WO1996012881A1 PCT/US1995/011818 US9511818W WO9612881A1 WO 1996012881 A1 WO1996012881 A1 WO 1996012881A1 US 9511818 W US9511818 W US 9511818W WO 9612881 A1 WO9612881 A1 WO 9612881A1
Authority
WO
WIPO (PCT)
Prior art keywords
translating sleeve
gas turbine
translating
turbine engine
downstream
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.)
Ceased
Application number
PCT/US1995/011818
Other languages
English (en)
Inventor
Paul W. Duesler
Constantino V. Lofredo
Harold T. Prosser, Jr.
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.)
RTX Corp
Original Assignee
United Technologies Corp
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 United Technologies Corp filed Critical United Technologies Corp
Publication of WO1996012881A1 publication Critical patent/WO1996012881A1/fr
Anticipated expiration legal-status Critical
Ceased legal-status Critical Current

Links

Classifications

    • 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
    • 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
    • F02K1/54—Nozzles having means for reversing jet thrust
    • F02K1/64—Reversing fan flow
    • F02K1/70—Reversing fan flow using thrust reverser flaps or doors mounted on the fan housing
    • F02K1/72—Reversing fan flow using thrust reverser flaps or doors mounted on the fan housing the aft end of the fan housing being movable to uncover openings in the fan housing for the reversed flow
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
    • F02K3/02—Plants 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
    • F02K3/04—Plants 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 the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
    • F02K3/06—Plants 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 the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type with front fan

Definitions

  • the present invention relates to a gas turbine engine and, more particularly, to a variable area fan exhaust nozzle therefor.
  • Important performance criteria for modern gas turbine engines include greater thrust, minimization of weight, and reduction in noise levels and fuel consumption. As is well known in the ail, a reduction in the fan pressure ratio improves the propulsive efficiency of a gas turbine engine.
  • the mass flow rate through the fan must be increased to maintain the same engine thrust.
  • Longer fan blades increase the mass flow rate.
  • reduction of the fan pressure ratio and an increase in the length of the fan blades adversely affect the fan stability.
  • Longer fan blades rotating at lower speeds pump additional air through the fan.
  • the additional mass flow at the lower fan pressure ratio contributes to the engine thrust as the air exits through a fan exhaust nozzle disposed downstream of the fan.
  • the additional air is restricted through the fan exhaust nozzle and the resulting back pressure on the fan negatively affects the aerodynamic stability of the fan.
  • fan stability is a limiting factor to low fan pressure ratio engines. Varying the pitch of the fan blades is one approach to control fan stability.
  • the pitch of the fan blades changes to tailor the amount of airflow passing through the fan during the different modes of operation of the gas turbine engine. During takeoff, climb and descent, the amount of air pumped by the fan blades is reduced, thereby reducing back pressure and avoiding instability conditions.
  • a gas turbine engine having a core engine enclosed in a conical core cowl and an outer nacelle with the outer nacelle being disposed radially outward of the core cowl and spaced apart therefrom, includes a translating sleeve disposed in the downstream portion of the outer nacelle to increase the effective area of the fan exhaust nozzle as the translating sleeve translates axially downstream to cooperate with the conical core cowl having a decreasing downstream diameter.
  • the fan exhaust nozzle is defined between the trailing edge of the translating sleeve and the core cowl.
  • the translating sleeve comprises an aerodynamically shaped body and a plurality of actuating means moving the translating sleeve from a fully stowed position axially downstream into a fully deployed position during climb, takeoff, and descent.
  • the translating sleeve is also capable of having a plurality of intermediate deployed positions.
  • variable area fan exhaust nozzle allows gas turbine engines to have higher efficiency at cruise without adversely effecting fan stability at other modes of operation.
  • the additional fan airflow improves the propulsive efficiency as the air exits through the fan exhaust nozzle.
  • the translating sleeve is translated axially downstream into the deployed position so that the effective area of the fan exhaust nozzle, defined between the trailing edge of the axially extended translating sleeve and reduced diameter core cowl, is increased.
  • the present invention improves the fuel consumption at cruise and reduces noise levels at takeoff, climb, and approach.
  • the plurality of intermediate deployed positions of the translating sleeve result in a gradual and continuous variation of area of the fan exhaust nozzle, allowing further optimization of performance of the gas turbine engine by improving weight of the overall engine and fuel consumption.
  • the translating sleeve comprises two semi-cylinders mating with each other to provide a continuous inner surface of the nozzle to withstand internal pressure and to minimize the leakage of airflow.
  • a major advantage of the present invention is that the translating sleeve is relatively simple structurally and is able to withstand "hoop" loading with relatively light weight structure.
  • FIG. 1 is a simplified, cross sectioned elevation of a gas turbine engine and a nacelle with a thrust reverser and a translating sleeve, according to the present invention
  • FIG. 2 is an enlarged, sectioned elevation of the thrust reverser and the translating sleeve of FIG. 1 at cruise with the thrust reverser and translating sleeve being depicted in a stowed position
  • FIG. 3 is a sectioned elevation of the thrust reverser and the translating sleeve of FIG. 2, at takeoff, climb, and descent with the translating sleeve being depicted in deployed position;
  • FIG. 4 is a sectioned elevation of the thrust reverser and the translating sleeve of FIG. 3 at reverse with both the thrust reverser and the translating sleeve being depicted in deployed position;
  • FIG. 5 is a diagrammatic, cross sectioned elevation of the thrust reverser and the translating sleeve of FIG. 4, taken along line 5-5. Best Mode for Carrying Out the Invention
  • a gas turbine engine 10 having a core engine 12 with a fan 14 disposed about a center longitudinal axis 16 includes an annular nacelle 18 encasing the core engine 12.
  • the annular nacelle 18 comprises an outer nacelle 20 having an upstream portion 22 and a downstream portion 24 and a conical core cowl 26 disposed radially inward from the outer nacelle 20 and spaced apart therefrom.
  • the outer nacelle 20 and the core cowl 26 define an annular flow path 28 therebetween.
  • a fan exhaust nozzle 30 is defined between a trailing edge 32 of the outer nacelle 20 and the core cowl 26.
  • the downstream portion 24 of the outer nacelle 20 includes a thrust reverser mechanism 36 and a variable area fan exhaust nozzle translating sleeve 38.
  • the thrust reverser mechanism 36 is of a conventional type, having a thrust reverser blocker door 50 and a thrust reverser movable body 52 with a recess 54 to accommodate a plurality of turning vanes 56 and a plurality of thrust reverser actuators 58 (superseded in FIG. 2) therein.
  • the turning vanes 56 that include a plurality of guide vanes 60, are secured onto a torque box 62 on the upstream end thereof and a support ring 64 on the downstream end thereof.
  • the thrust reverser actuators 58 are hydraulic actuators of the conventional type, having a cylinder 66 and a moveable rod 68, wherein the cylinder is secured onto the torque box 62 and the rod 68 is secured onto an inner surface of the recess 54. Hydraulic pressure to the actuators is provided through tubing 70.
  • the thrust reverser blocker door 50 is disposed radially inward of the thrust reverser body 52 and is in a substantially parallel relationship to the longitudinal axis 16 when in a stowed position, as shown in FIGs. 1 and 2.
  • the thrust reverser blocker door 50 pivots about a pivot point 71 into a deployed position, as shown in FIG. 4.
  • the fan exhaust nozzle translating sleeve 38 having an aerodynamically shaped outer surface 72 and inner surface 74, is disposed radially inward of the thrust reverser 36 and radially outward of the blocker door 50 and extends downstream of the thrust reverser 36.
  • a plurality of translating sleeve actuators 76 provide axial translation to the translating sleeve 38.
  • Each actuator 76 is of a hydraulic type, having a cylinder 78 and a moveable rod 80, with the cylinder 78 being secured onto the torque box 62 and the rod 80 being secured onto the translating sleeve 38.
  • An aerodynamic flap seal 82 is fixedly attached to the most downstream portion of the thrust reverser 36 to bridge the gap between the thrust reverser 36 and the translating sleeve 38 to ensure an aerodynamically smooth exterior surface of the outer nacelle 20.
  • An inflatable seal 84 is disposed between the translating sleeve 38 and the thrust reverser 36 to reduce air leakage therebetween during translation. Alternatively, a lip seal or any other type of a seal may be used to prevent air leakage.
  • a translating sleeve bumper seal 85 is disposed on a leading edge of the translating sleeve 38 and bears against the pivot point 71 , when the translating sleeve 38 is in a fully stowed position.
  • a thrust reverser bumper seal 86 is disposed on the leading edge of the inner wall of the thrust reverser body 52 and bears against the torque box 62 when the thrust reverser 36 is in the fully stowed position to reduce air leakage.
  • the thrust reverser body 52 and the translating sleeve 38 each comprises two semi-cylinders 88, 89 and 90, 91 , respectively.
  • Each semi-cylinder 90, 91 of the translating sleeve 38 includes longitudinal edges 92-93, 94-95, respectively.
  • Each longitudinal edge 92-95 has a T-slider 96 attached thereto.
  • the T-sliders 96 of the longitudinal edges 92, 94 slidingly engage tracks 97 of a hinge mechanism 98.
  • the hinge mechanism 98 includes a mounting hinge 99 securing the hinge mechanism 98 onto a pylon (not shown), which is attached to the wing of the airplane (not shown).
  • the T-sliders 96 of the longitudinal edges 92-95 of the translating sleeve 38 slidingly engage tracks 97 of a latch mechanism 100, which is disposed substantially diagonally across from the hinge mechanism 98.
  • the semi-cylinders 90, 91 of the translating sleeve 38 and the latch and hinge mechanisms 100, 98 form a substantially continuous annulus with the substantially continuous surface 74.
  • the thrust reverser 36 has a mounting structure analogous to that of the translating sleeve 38.
  • the latch mechanism 100 can be opened to allow the two sets of semi-cylinders 88- 89, 90-91 of the thrust reverser 36 and of the translating sleeve 38 to pivot about the hinge mechanism 98 to allow access to the core engine 12 disposed therein.
  • O-ring seals (not shown) are disposed at the ends of the tracks 97 to prevent air leakage between the tracks 97 and the T-sliders 96. In cruise mode, shown in FIGS.
  • both the thrust reverser 36 and the fan exhaust nozzle translating sleeve 38 are in a fully stowed position with the movable rods 68, 80 of the thrust reverser actuators 58 and of the translating sleeve actuators 76 in their retracted positions.
  • the translating sleeve 38 is fully deployed.
  • the hydraulic pressure activates the plurality of translating sleeve actuators 76 so that the moveable rods 80 extend axially downstream, thereby transmitting axial, downstream movement to the translating sleeve 38.
  • the sliding downstream movement is effectuated as the T-sliders 96 disposed along the longitudinal edges 92- 95 of the semi-cylinders 90, 91 of the translating sleeve 38 slidingly engage the tracks 97 of the hinge and latch mechanisms 98, 100.
  • the annular area of the effective fan exhaust nozzle 30' defined by the trailing edge 32 of the translating sleeve 38 and the core cowl 26 is increased.
  • the effective increase in the annular area of the fan nozzle is gained due to the decreasing downstream diameter of the conical core cowl 26.
  • the greater area fan nozzle 30' becomes the controlling area for the exiting airflow 28 at takeoff, climb, and descent.
  • the aerodynamically shaped inner surface 74 of the translating sleeve 38 insures that the airflow is not choked upstream of the fan exhaust nozzle 30'.
  • a number of intermediate deployed positions between the fully stowed position and fully deployed position for the translating sleeve are possible with the hydraulic actuators being activated gradually.
  • the thrust reverser 36 may be activated to provide reverse thrust to the gas turbine engine 10.
  • the translating sleeve 38 must also be activated to expose the turning vanes 56 to the airflow 28.
  • the translating sleeve 38 is activated and translates downstream in the same manner as described above.
  • the thrust reverser 36 is moved axially downstream when the hydraulic pressure builds up in the thrust reverser cylinders 66 and extends the moveable rods 68 axially downstream.
  • the thrust reverser 36 then slidingly moves downstream as the thrust reverser T-slides 96 slide downstream in the tracks 97.
  • the reverser door 50 pivots radially inward to block the air flow path 28 from exiting through the fan exhaust nozzle 30', thereby redirecting the airflow to pass through the guide vanes 60.
  • the gas turbine engines that employ the present invention can achieve higher propulsive efficiency with a lower pressure ratio and higher mass flow without sacrificing engine thrust and without fan stability problems. At cruise, as the translating sleeve is in the fully stowed position, the gas turbine engines with lower pressure ratios enjoy higher thrust, reduced noise levels and improved fuel consumption.
  • the increased area of the fan exhaust nozzle 30' allows the additional air flow generated by the fan to exit the engine 10 without causing excessive back pressure on the fan blades and thus without stalling the fan 14.
  • gradually and continuously varying area of the fan exhaust nozzle allows further optimization of the performance by reducing the overall weight of the gas turbine engine and improving fuel consumption.
  • the combination of the lower fan pressure ratio and the higher fan mass flow reduces noise levels at approach, takeoff and climb.
  • the noise reduction is a result of two factors. First, the fan rotating at lower speeds thereby producing less noise. Second, the extended downstream translating sleeve affords an additional attenuation path that reduces noise levels. Moreover, the combination of the lower fan pressure ratio and the higher fan mass flow improves fuel consumption at cruise.
  • the variable area fan exhaust nozzle having the translating sleeve 38 of the present invention allows an increase in the area of the nozzle 30' in excess of 40%.
  • the translating sleeve 38 achieves a significant increase in the area of the nozzle 30' without a significant weight penalty to the gas turbine engine 10.
  • the two semi-cylinders 90, 91 of the translating sleeve 38 form a continuous inner surface 74 of the fan exhaust nozzle.
  • the continuous surface 74 allows the nozzle to withstand the internal pressure of the airflow without excessive weight and to avoid air leakage.
  • this invention can be used with a thrust reverser having translating turning vanes, rather than stationary turning vanes 56, as described in the preferred embodiment.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Control Of Turbines (AREA)

Abstract

Cette invention se rapporte à un moteur à turbine à gaz (10) qui comprend un manchon de translation (38) disposé dans une partie aval d'une nacelle externe (20). Une buse (30) d'évacuation de l'air de ventilation, à section de superficie variable, est formée entre le bord de fuite du manchon de translation (38) et un pavillon à âme conique (26) disposé radialement vers l'intérieur de la nacelle externe (20) et à une certaine distance de celle-ci. Le manchon de translation (38) effectue un mouvement de translation vers l'aval, afin de coopérer avec le diamètre décroissant du pavillon à âme conique (26), pour que la superficie de la section de la buse (30) d'évacuation de l'air de ventilation augmente.
PCT/US1995/011818 1994-10-20 1995-09-15 Buse d'evacuation d'air de ventilation, a section de superficie variable Ceased WO1996012881A1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US32662194A 1994-10-20 1994-10-20
US326,621 1994-10-20

Publications (1)

Publication Number Publication Date
WO1996012881A1 true WO1996012881A1 (fr) 1996-05-02

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Cited By (14)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0779429A3 (fr) * 1995-12-14 1999-04-21 United Technologies Corporation Tuyère à section variable pour turbosoufflante
FR2830051A1 (fr) * 2001-09-27 2003-03-28 Hurel Hispano Le Havre Systeme de verrouillage sur un inverseur de poussee a grilles
GB2443947A (en) * 2006-11-14 2008-05-21 Gen Electric Varying turbofan engine nozzle throat area
GB2443948A (en) * 2006-11-14 2008-05-21 Gen Electric Turbofan engine baffle assembly for varying throat of fan nozzle duct
FR2929655A1 (fr) * 2008-04-02 2009-10-09 Aircelle Sa Nacelle de turboreacteur a double flux
WO2011014346A3 (fr) * 2009-07-31 2011-04-07 General Electric Company Ensemble inverseur de poussée intégré
FR2971015A1 (fr) * 2011-02-01 2012-08-03 Snecma Tuyere d'ejection pour turboreacteur d'avion a double flux separes a capot secondaire deployable et corps central retractable
EP2466101A3 (fr) * 2010-12-15 2015-02-18 GE Aviation Systems LLC Système et procédé de fonctionnement d'un inverseur de poussée pour système de propulsion de réacteur à double flux
EP2989313A4 (fr) * 2013-04-24 2016-04-20 United Technologies Corp Moteur de turbine à engrenages comportant un conduit en o et un inverseur de poussée
EP2964944A4 (fr) * 2013-03-04 2016-11-16 United Technologies Corp Inverseur de poussée à portes pivotantes comportant une buse de section variable
US20170016413A1 (en) * 2015-07-13 2017-01-19 The Boeing Company Telescoping electrical cable
US10006405B2 (en) 2012-11-30 2018-06-26 General Electric Company Thrust reverser system with translating-rotating blocker doors and method of operation
EP3772581A1 (fr) * 2019-08-05 2021-02-10 Rohr, Inc. Système d'entraînement pour structure de translation
US11313322B2 (en) 2018-07-19 2022-04-26 Rolls-Royce Plc Exhaust nozzle assembly

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3779010A (en) * 1972-08-17 1973-12-18 Gen Electric Combined thrust reversing and throat varying mechanism for a gas turbine engine
US3797785A (en) * 1972-08-31 1974-03-19 Rohr Industries Inc Thrust modulating apparatus
EP0191204A1 (fr) * 1983-05-19 1986-08-20 The Boeing Company Elément de glissière démontable pour un carénage mobile en translation d'une nacelle d'un moteur à réaction
US4802629A (en) * 1982-10-22 1989-02-07 The Boeing Company Plug-type exhaust nozzle having a variable centerbody and translating shroud

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3779010A (en) * 1972-08-17 1973-12-18 Gen Electric Combined thrust reversing and throat varying mechanism for a gas turbine engine
US3797785A (en) * 1972-08-31 1974-03-19 Rohr Industries Inc Thrust modulating apparatus
US4802629A (en) * 1982-10-22 1989-02-07 The Boeing Company Plug-type exhaust nozzle having a variable centerbody and translating shroud
EP0191204A1 (fr) * 1983-05-19 1986-08-20 The Boeing Company Elément de glissière démontable pour un carénage mobile en translation d'une nacelle d'un moteur à réaction

Cited By (28)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0779429A3 (fr) * 1995-12-14 1999-04-21 United Technologies Corporation Tuyère à section variable pour turbosoufflante
FR2830051A1 (fr) * 2001-09-27 2003-03-28 Hurel Hispano Le Havre Systeme de verrouillage sur un inverseur de poussee a grilles
EP1298309A1 (fr) * 2001-09-27 2003-04-02 Hurel-Hispano Le Havre Système de verrouillage sur un inverseur de poussée à grilles
WO2003027474A1 (fr) * 2001-09-27 2003-04-03 Hurel-Hispano Le Havre Systeme de verrouillage sur un inverseur de poussee a grilles
US8091334B2 (en) 2006-11-14 2012-01-10 General Electric Company Method of operating a turbofan engine cowl assembly
GB2443947A (en) * 2006-11-14 2008-05-21 Gen Electric Varying turbofan engine nozzle throat area
GB2443948A (en) * 2006-11-14 2008-05-21 Gen Electric Turbofan engine baffle assembly for varying throat of fan nozzle duct
US7886518B2 (en) 2006-11-14 2011-02-15 General Electric Company Turbofan engine cowl assembly and method of operating the same
GB2443947B (en) * 2006-11-14 2011-09-07 Gen Electric Turbofan engine nozzle assembly and method for operating the same
GB2443948B (en) * 2006-11-14 2011-09-07 Gen Electric Turbofan engine cowl assembly and method of operating the same
FR2929655A1 (fr) * 2008-04-02 2009-10-09 Aircelle Sa Nacelle de turboreacteur a double flux
WO2009125157A3 (fr) * 2008-04-02 2009-12-03 Aircelle Nacelle de turboréacteur à double flux
US8713910B2 (en) 2009-07-31 2014-05-06 General Electric Company Integrated thrust reverser/pylon assembly
WO2011014346A3 (fr) * 2009-07-31 2011-04-07 General Electric Company Ensemble inverseur de poussée intégré
EP2466101A3 (fr) * 2010-12-15 2015-02-18 GE Aviation Systems LLC Système et procédé de fonctionnement d'un inverseur de poussée pour système de propulsion de réacteur à double flux
FR2971015A1 (fr) * 2011-02-01 2012-08-03 Snecma Tuyere d'ejection pour turboreacteur d'avion a double flux separes a capot secondaire deployable et corps central retractable
US10006405B2 (en) 2012-11-30 2018-06-26 General Electric Company Thrust reverser system with translating-rotating blocker doors and method of operation
EP2964944A4 (fr) * 2013-03-04 2016-11-16 United Technologies Corp Inverseur de poussée à portes pivotantes comportant une buse de section variable
US10400621B2 (en) 2013-03-04 2019-09-03 United Technologies Corporation Pivot door thrust reverser with variable area nozzle
EP2989313A4 (fr) * 2013-04-24 2016-04-20 United Technologies Corp Moteur de turbine à engrenages comportant un conduit en o et un inverseur de poussée
US20170016413A1 (en) * 2015-07-13 2017-01-19 The Boeing Company Telescoping electrical cable
US10422301B2 (en) * 2015-07-13 2019-09-24 The Boeing Company Telescoping electrical cable
US11313322B2 (en) 2018-07-19 2022-04-26 Rolls-Royce Plc Exhaust nozzle assembly
EP3772581A1 (fr) * 2019-08-05 2021-02-10 Rohr, Inc. Système d'entraînement pour structure de translation
US11391242B2 (en) 2019-08-05 2022-07-19 Rohr, Inc. Drive system for translating structure
US11391241B2 (en) 2019-08-05 2022-07-19 Rohr, Inc. Drive system for translating structure
EP4276299A3 (fr) * 2019-08-05 2024-01-24 Rohr, Inc. Système d'entraînement pour structure de translation
US12031500B2 (en) 2019-08-05 2024-07-09 Rohr Inc. Drive system for translating structure

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