Stabilizer with structural box and sacrificial surfaces

Abstract

An aircraft includes a tail extending from a fuselage. The tail defines a structural box having first and second vertical stabilizers that support a horizontal stabilizer. The tail includes at least one sacrificial control surface and at least one primary control surface. The primary control surfaces maintain aircraft controllability in the event that the sacrificial control surface becomes inoperable.

Claims

What is claimed is: 1. An aircraft comprising; a fuselage; and a tail extending from the fuselage, the tail defining a structural box having first and second vertical stabilizers and a horizontal stabilizer supported atop the first and second vertical stabilizers, the tail including at least one sacrificial control surface and at least one primary control surface, the at least one sacrificial control surface including a defined frangible connection along which the sacrificial control surface breaks away from the tail responsive to a strike from debris, wherein the primary control surface maintains aircraft controllability upon the at least one sacrificial control surface breaking away from the tail. 2. The aircraft as recited in claim 1 , wherein the at least one sacrificial control surface and primary control surface are part of the horizontal stabilizer. 3. The aircraft as recited in claim 2 , wherein the horizontal stabilizer is spaced apart from the fuselage in a direction that is substantially perpendicular to a longitudinal centerline of the fuselage. 4. The aircraft recited in claim 2 , wherein the at least one primary control surface is supported across the first and second vertical stabilizers. 5. The aircraft recited in claim 4 , wherein the structural box includes the first and second vertical stabilizers and the at least one primary control surface that is supported across the first and second vertical stabilizers. 6. The aircraft as recited in claim 1 , including a propulsion system mounted to an aft end of the fuselage and a burst zone defined about the propulsion system that encompasses the at least one sacrificial control surface. 7. The aircraft recited in claim 6 , wherein the propulsion system comprises a gas turbine engine including a gas generator disposed about a first axis driving a fan section disposed about a second axis angled relative to the first axis. 8. The aircraft recited in claim 7 , wherein the first axis is angled relative to a longitudinal centerline of the fuselage. 9. The aircraft recited in claim 7 , wherein the gas generator includes a fan drive gear system for driving the fan section. 10. An aircraft assembly comprising a fuselage including a forward portion and an aft portion; a turbine engine mounted within the aft portion, wherein a burst zone is defined about the turbine engine; and a tail disposed at least partially with the burst zone, the tail portion including a horizontal stabilizer supported across a first vertical stabilizer and a second vertical stabilizer, wherein the horizontal stabilizer further includes at least one sacrificial control surface including a defined frangible connection to the horizontal stabilizer along which the at least one sacrificial control surface breaks away from the horizontal stabilizer in a defined manner responsive to a strike to the sacrificial control surface by a foreign object within the defined burst zone. 11. The aircraft assembly as recited in claim 10 , wherein a primary control surface extends across the first vertical stabilizer and the second vertical stabilizer. 12. The aircraft assembly as recited in claim 10 , wherein the at least one sacrificial control surface is disposed at distal ends of the horizontal stabilizer. 13. The aircraft assembly as recited in claim 12 , wherein the primary control surface is disposed between sacrificial control surfaces. 14. The aircraft assembly as recited in claim 10 , wherein the turbine engine comprises first and second turbine engines defining corresponding first and second burst zones. 15. The aircraft assembly as recited in claim 14 , wherein the first and second turbine engines include corresponding gas generator sections disposed about different axes that are angled away from each other. 16. The aircraft assembly as recited in claim 15 , wherein the gas generators are angled relative to a longitudinal centerline of the fuselage.
CROSS REFERENCE TO RELATED APPLICATION This application claims priority to U.S. Provisional Application No. 61/726,737 filed Nov. 15, 2012 and U.S. Provisional Application No. 61/735,717 filed Dec. 11, 2012. BACKGROUND Conventional aircraft architecture includes wing mounted gas turbine engines. In some aircraft architectures gas turbine engines are mounted atop the fuselage or on opposite sides of the aircraft fuselage. Commercial aircraft typically utilize gas turbine engines that in include a fan section driven by a core engine or gas generator. The gas generator engine includes a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustion section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section through a driven shaft. Alternate aircraft architectures may require alternate mounting locations of the gas turbine engines to enable specific wing and fuselage configurations. However, conventional gas turbine engine configurations have been developed to operate with conventional aircraft architectures. Accordingly, alternate gas turbine engine configurations may be required to enable implementation of favorable aspects of alternate aircraft architectures. SUMMARY An aircraft fuselage, according to an exemplary embodiment of this disclosure includes a tail extending from a fuselage. The tail defines a structural box that has first and second vertical stabilizers that support a horizontal stabilizer. The tail includes at least one sacrificial control surface and at least one primary control surfaces. The primary control surfaces maintain aircraft controllability separate from the at least one sacrificial control surface. In a further embodiment of the above, the sacrificial control surface and primary control surfaces are part of the horizontal stabilizer. In a further embodiment of the above, the horizontal stabilizer is spaced apart from the fuselage in a direction that is substantially perpendicular to a longitudinal centerline of the fuselage. In a further embodiment of the above, the primary control surface is disposed between the first and second vertical stabilizers. In a further embodiment of the above, the structural box includes the first and second vertical stabilizers and the primary control surfaces disposed between the first and second vertical stabilizers. In a further embodiment of the above, a propulsion system is mounted to an aft end of the fuselage. A burst zone is defined about the propulsion system that encompasses at least one sacrificial control surface. In a further embodiment of the above, the propulsion system includes a gas turbine engine with a gas generator that is disposed about a first axis that drives a fan section that is disposed about a second axis angled relative to the first axis. In a further embodiment of the above, the first axis is angled relative to a longitudinal centerline of the fuselage. In a further embodiment of the above, the gas generator includes a fan drive gear system for driving the fan section. An aircraft assembly according to another exemplary embodiment includes a fuselage having a forward portion and an aft portion and a turbine engine mounted within the aft portion. A burst zone is defined about the turbine engine. A tail is disposed at least partially with the burst zone. The tail portion has a horizontal stabilizer supported across a first vertical stabilizer and a second vertical stabilizer. The horizontal stabilizer further includes at least one sacrificial control surface within the defined burst zone that is frangible from the horizontal stabilizer. In a further embodiment of the above, a primary control surface is disposed between the first vertical stabilizer and the second vertical stabilizer. In a further embodiment of the above, the sacrificial control surfaces are disposed at distal ends of the horizontal stabilizer. In a further embodiment of the above, the primary control surface is disposed between sacrificial control surfaces. In a further embodiment of the above, the turbine engine comprises first and second turbine engines defining corresponding first and second burst zones. In a further embodiment of the above, the first and second turbine engines include corresponding gas generator sections disposed about different axes that are angled away from each other. In a further embodiment of the above, the gas generators are angled relative to a longitudinal centerline of the fuselage. Although the different examples have the specific components shown in the illustrations, embodiments of this disclosure are not limited to those particular combinations. It is possible to use some of the components or features from one of the examples in combination with features or components from another one of the examples. These and other features disclosed herein can be best understood from the following specification and drawings, the following of which is a brief description. BRIEF DESCRIPTION OF THE DRAWINGS FIG. 1 is a schematic view of an example aircraft including an aft mounted propulsion system. FIG. 2 is a schematic view of an example propulsion system. FIG. 3 is a schematic view of an example burst zone. FIG. 4 is a top view of an example tail assembly. FIG. 5 is a schematic view of a separated portion of the example tail assembly. FIG. 6 is a top view of the example tail assembly with a separated sacrificial secondary control surface. DETAILED DESCRIPTION Referring to the FIGS. 1 and 2 an aircraft 10 includes a fuselage 12 having wings 16 and a tail 14 . A propulsion system 18 is mounted aft end of the fuselage 12 . The propulsion system 18 includes first and second gas turbine engines. The gas turbine engines include first and second gas generators 20 a - b that drives corresponding first and second fan sections 22 a - b. Each of the first and second gas generators 20 a - b are disposed about an engine axis A and drive the corresponding fan sections 22 a - b about a second axis B. The first axis A and second axis B are angled relative to each other. In traditional engine architectures, the axis of the gas generator is aligned in the same direction as the axis of the propulsor (or fan). When the gas generator is rotated more than 90 degrees relative to the propulsor, it is considered a reverse core engine. This configuration allows the fan to be driven by a free turbine 25 a - b , which is powered by the exhaust from the gas generator. The free turbine 25 a - b may drive a fan drive gear system 27 a - b that enables the free turbine 25 a - b and fan to rotate at different rotational speeds. The gas generators in this example include a compressor 24 , a combustor 26 and a turbine 28 . Air is drawn in through inlets 32 a - b to the compressor 24 is compressed and communicated to a combustor 26 . In the combustor 26 , air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through the turbine 28 where some energy is extracted and utilized to drive the compressor 24 . The output from the gas generator is a stream of high-pressured gas that drives the free turbine, and its corresponding fan 22 a - b. In the disclosed example, each of the first and second fans 22 a - b is mounted substantially parallel to each other about respective second axes A. The two axes A are also disposed substantially parallel to aircraft longitudinal axis C. Gas turbine engines are not typically mounted next to each other due to practical limitations related to overall aircraft survivability in the event of engine failure. A burst zone is defined around gas turbine engines within which another gas turbine engine is not permitted due to possible fragmentation from one failed engine disabling the second engine. The disclosed gas generators 20 a - b are disposed along a second axis B at an angle 30 relative to the corresponding axes A and to each other such that neither gas generator 20 a - b is disposed within a burst zone 34 of the other gas generator 20 a - b . Each of the gas generators 20 a - b is disposed at an angle 50 away from the other gas generator 20 a - b such that each is orientated outside of the others bust zone 34 . The gas generators 20 a - b are further set at an angle 52 relative to the aircraft longitudinal axis C. Referring to FIG. 3 with continued reference to FIGS. 1 and 2 , the aircraft tail 14 includes a first vertical stabilizer 38 and a second vertical stabilizer 40 that support a horizontal stabilizer 36 . The horizontal stabilizer 36 extends across the first and second vertical stabilizers 38 , 40 and includes a primary control surface 44 and secondary control surfaces 46 a - b . The primary control surface 44 along with the first and second vertical stabilizers 38 , 40 define a structural box 42 that is at least partially disposed within the burst zone 34 . In the disclosed aircraft architecture, portions of the horizontal stabilizer 36 are within the burst zone 34 defined by the angled orientation of the gas generators 22 a - b . In this example, the secondary control surfaces 46 a - b are disposed within the burst zone 34 . The secondary control surfaces 46 a - b define regions within the burst zone that are sacrificial surfaces designed to break away in a controlled manner such that aircraft control is maintained. The horizontal stabilizer 36 includes frangible connections 48 that break away in a controlled manner to enable the aircraft 10 to maintain stability and control. The non-break away surfaces are part of the structural box 42 and include the primary control surface 44 that maintains the desired aircraft control after loss of a secondary control surface 46 a - b. The example vertical stabilizers 38 , 40 define the structural box 42 that is resistant to damage from potential fragments within the bust zone. The structural box 42 includes portions of the fuselage 12 , the first and second vertical stabilizers 38 , 40 and the primary control surface 44 of the horizontal stabilizer 36 . The structural box 42 is strengthened relative to the surrounding structures to provide a level of survivability desired to maintain the primary control surface 44 of the horizontal stabilizer 36 . Referring to FIGS. 4, 5 and 6 , the tail 14 is shown with the frangible connections 48 as part of the horizontal stabilizer 36 . The frangible connections 48 are disposed on either side of the primary control surfaces 44 . The primary control surface 44 is supported between the first and second vertical stabilizers 48 , 40 . The secondary control surfaces 46 a - b that define the sacrificial surfaces are disposed on distal ends of the horizontal stabilizer 36 within the burst zones 34 . In operation, during a fragmentation event of one of the gas generators 20 a - b , fragments may be present within the burst zone 34 and result in damage to one of the sacrificial secondary control surface 46 a as is schematically shown in FIGS. 5 and 6 . The sacrificial secondary control surface 46 a may sustain damage without separating from the tail 14 and/or may separate in a controlled manner along the frangible connection 48 . In either instance, the primary control surface 44 remains intact to provide aircraft stability and control. Accordingly, the example aircraft architecture includes features that enable the use and operation of control surfaces within the burst zones by including controlled break away portions in addition to a structural box outside of the burst zone to maintain integrity of control surfaces outside of the burst zones. Although an example embodiment has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this disclosure. For that reason, the following claims should be studied to determine the scope and content of this disclosure.

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Patent Citations (45)

    Publication numberPublication dateAssigneeTitle
    GB-526775-ASeptember 25, 1940Paul Aime RichardImprovements in or relating to aircraft-stabilizing devices
    JP-2008075582-AApril 03, 2008Japan Aerospace Exploration Agency, 独立行政法人 宇宙航空研究開発機構V/stol機用推進システム
    US-2004025493-A1February 12, 2004Wojciechowski Paul M.High bypass multi-fan engine
    US-2004245374-A1December 09, 2004Morgan Brian H.Vertical take-off and landing aircraft
    US-2005262682-A1December 01, 2005Fatigue Technology, Inc.Method and apparatus employing eccentric bushing
    US-2006144991-A1July 06, 2006Aldo FredianiSwept-wing box-type aircraft with high fligh static stability
    US-2008073459-A1March 27, 2008Airbus FranceMulti-Engine Aircraft
    US-2009020643-A1January 22, 2009Airbus FranceAircraft having reduced environmental impact
    US-2009065632-A1March 12, 2009Airbus FranceAircraft with jet engines arranged at the rear
    US-2009084889-A1April 02, 2009Airbus FranceAircraft having a reduced acoustic signature
    US-2010032520-A1February 11, 2010Airbus Uk LimitedControl surface failsafe drop link
    US-2010059623-A1March 11, 2010Airubus OperationsAircraft with its fuselage suspended under the wing
    US-2010133377-A1June 03, 2010Airbus Operations (Societe Par Actions Simplifiee)Airplane with flat rear fuselage said queue-de-morue empennage
    US-2010148000-A1June 17, 2010Airbus Espana S.L..Aircraft horizontal stabilizer surface
    US-2010212288-A1August 26, 2010Williams International Co., L.L.C.Jet Engine Exhaust Nozzle and Associated System and Method of Use
    US-2010243810-A1September 30, 2010Airbus Espana, S.L.Integrative structure for aircraft fairing
    US-2010264264-A1October 21, 2010Airbus FranceMethod for producing an aircraft with reduced envrionmental impact and the aircraft thus obtained
    US-2011036939-A1February 17, 2011William Craig EasterRapidly convertible hybrid aircraft and manufacturing method
    US-2011061579-A1March 17, 2011Van Gelder Klaas BoudewijnDynamic fin comprising coupled fin sections
    US-2011220758-A1September 15, 2011Airbus FranceArchitecture d'avion a fuselage large
    US-2012104184-A1May 03, 2012Airbus Operations Sas, Airbus Operations GmbhAircraft comprising a device for influencing the directional stability of the aircraft, and a method for influencing the directional stability of the aircraft
    US-2012168559-A1July 05, 2012Bell Helicopter Textron Inc.Removable Horizontal Stabilizer for Helicopter
    US-2012272656-A1November 01, 2012United Technologies CorporationMultiple core variable cycle gas turbine engine and method of operation
    US-2013001356-A1January 03, 2013Airbus Operations, S.L.Reinforced aircraft fuselage
    US-3972490-AAugust 03, 1976Mcdonnell Douglas CorporationTrifan powered VSTOL aircraft
    US-4291853-ASeptember 29, 1981The Boeing CompanyAirplane all-moving airfoil with moment reducing apex
    US-4447022-AMay 08, 1984Lion Charles EReduced noise monolithic wing-stabilizer aircraft
    US-4448372-AMay 15, 1984The Boeing CompanyAircraft vertical fin-fuselage structural integration system
    US-4500055-AFebruary 19, 1985Dornier GmbhAircraft propulsion system arrangement
    US-5131605-AJuly 21, 1992Grumman Aerospace CorporationFour engine VTOL aircraft
    US-5289996-AMarch 01, 1994The United States Of America As Represented By The Secretary Of The Air ForceAircraft windshield system with frangible panel for aircrew emergency escape
    US-5445346-AAugust 29, 1995Gilbert; Raymond D.Aircraft tail surface spoilers
    US-5979824-ANovember 09, 1999Gagliano; Christopher, Boadman; Thomas E.Stabilizer fins-inverted for aircraft
    US-6273363-B1August 14, 2001Daimlerchrysler Aerospace Airbus GmbhAircraft with a double-T tail assembly
    US-6543718-B2April 08, 2003Rolls-Royce PlcEngine arrangement
    US-6792746-B2September 21, 2004National Aerospace Laboratory Of JapanSeparated core engine type turbofan engine
    US-6824092-B1November 30, 2004Supersonic Aerospace International, LlcAircraft tail configuration for sonic boom reduction
    US-7107755-B2September 19, 2006Rolls-Royce PlcEngine arrangement
    US-7540450-B2June 02, 2009Pratt & Whitney Canada Corp.Aircraft propulsion system
    US-7600717-B2October 13, 2009Airbus Operations LimitedAircraft wings and fuel tanks
    US-7775834-B2August 17, 2010Lockheed Martin CorporationRugged, removable, electronic device
    US-7780116-B2August 24, 2010EurocopterComposite anti-crash structure with controlled buckling for an aircraft
    US-8015796-B2September 13, 2011United Technologies CorporationGas turbine engine with dual fans driven about a central core axis
    US-8167239-B2May 01, 2012Airbus Operations SasTurbojet for aircraft, aircraft equipped with such a turbojet, and method for mounting such a turbojet on an aircraft
    US-8186617-B2May 29, 2012Airbus Operations S.L.Aircraft having a lambda-box wing configuration

NO-Patent Citations (3)

    Title
    European Search Report for EP Application No. 13856025.5 dated Oct. 26, 2015.
    International Preliminary Report on Patentability for International Application No. PCT/US2013/065417 dated May 28, 2015.
    International Search Report and Written Opinion for International Application No. PCT/US2013/065417 dated Jan. 9, 2014.

Cited By (0)

    Publication numberPublication dateAssigneeTitle