Abstract:
The present invention discloses a novel fuel structure for housing and delivering disparate cryogenic fuels to combustion zones in an aerospace vehicle. The tank comprises a plurality of containers having volumes that are separated by common wall bulkheads and which are arranged substantially side-by-side in conformance with the interior of the aerospace vehicle. A tank support structure positioned within the vehicle interior includes lengthwise supports as well as cross-wise supports, with the latter including openings within which the rear ends of the containers are supported. Fuel from the containers is delivered to the vehicle&#39;s combustion system via appropriate fuel lines carried by dome shaped end caps at the rear ends of the containers.

Description:
The invention described herein was made in the performance of work under NASA Contract No. NCC8-115 and is subject to provisions of Section 305 of the National Aeronautics and Space Act of 1958 (42 U.S.C. 2457). 
     BACKGROUND OF THE INVENTION 
     1. Field of the Invention 
     The present invention relates to cryogenic propellant fuel tanks for air vehicles, and more particularly to a cryogenic propellant tank structure including multiple contiguous fuel containers having an overall configuration that conforms to the interior of the air vehicle, adjacent fuel containers sharing a common bulkhead structure. 
     2. Description of the Related Art 
     It is now well-known that in order to achieve a single stage to orbit reusable launch vehicle (“SSTO-RLV”), it is necessary to attain an extremely low vehicle mass fraction. “Mass fraction” is a term that commonly refers to the ratio of the dry weight of a vehicle to the gross lift-off weight of the same vehicle. 
     Various configurations of aerospace vehicles have been proposed to achieve such an objective. For example, the prior art teaches traditional SSTO configurations of the type typically used in the NASA&#39;s space shuttle program. These vehicles use external expendable drop fuel tanks, and examples are shown in U.S. Pat. No. 3,929,306 to Faget et al, U.S. Pat. No. 4,452,412 to von Pragenau, U.S. Pat. No. 4,557,444 to Jackson et al., and U.S. Pat. No. 4,817,890 to Coffinberry. SSTO vehicles which do not use expendable fuel tanks are also known, as evidenced by the teachings of U.S. Pat. No. 3,261,571 to Pinnes, U.S. Pat. No. 3,955,784 to Salkeld, and U.S. Pat. No. 5,975,466 to Kahara et al. In the Pinnes patent, the fuselage of an aircraft intended for orbital flight comprises a plurality of cryogenic fuel tanks arranged in a triangular configuration and secured together to form the fuselage of the craft. In the Kahara et al. patent, the aircraft disclosed includes a fuel tank, which takes the form of a collapsible bladder. And in the Salkeld patent, there is disclosed an aerospace vehicle which incorporates two different propulsion systems that are operated in sequence and which use cryogenic fuels stored in tanks within the body of the vehicle. 
     Against this background of known technology, the applicants have developed a novel fuel tank structure for housing and delivering disparate cryogenic fuels to combustion zones in an aerospace vehicle. The tank comprises a plurality of containers having volumes that are separated by common wall bulkheads and which are arranged substantially side-by-side in conformance with the interior of the aerospace vehicle. A tank support structure positioned within the vehicle interior includes lengthwise supports as well as cross-wise supports, with the latter including openings within which the rear ends of the containers are supported. Fuel from the containers is delivered to the vehicle&#39;s combustion system via appropriate fuel lines carried by dome shaped end caps at the rear ends of the containers. 
     OBJECTS AND SUMMARY OF THE INVENTION 
     It is therefore an object of the present invention to provide a novel cryogenic fuel tank for an aerospace vehicle that will provide an extremely low vehicle mass fraction, while overcoming many of the disadvantages and drawbacks of similar fuel tanks known in the art. 
     Another object of the present invention is to provide a fuel tank for cryogenic propellants, which conform to the interior space of an air vehicle in which the tank is located. 
     Still another object of the invention is to provide one fuel tank within another fuel tank, the two tanks holding disparate cryogenic fuels, which are intended to be mixed before being combusted. 
     Yet another object of the invention is to provide a novel cryogenic propellant tank in which the overall configuration of the tank is a deltoid shape. 
     These and other objects, advantages and features of the invention will become more apparent, as will equivalent structures which are intended to be covered herein, with the teaching of the principles of the invention in connection with the disclosure of the preferred embodiments thereof in the specification, claims and drawings in which: 
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS 
     FIG. 1 depicts the propellant tank  10  conformally disposed in an aerospace vehicle (the latter being shown in phantom lines); 
     FIG. 2 is a cross-sectional view of the aerospace vehicle shown in FIG. 1; 
     FIGS. 2A-2D depict the cross-sections of the aerospace vehicle shown in FIG. 1 taken along section lines A—A, B—B, C—C and D—D, respectively; 
     FIGS. 3-5 show structural framework for supporting the cryogenic propellant fuel tank in the aerospace vehicle; 
     FIG. 6 is a cross section through the rear portion of the fuel tank; 
     FIG. 7 is a Y-joint member depicted in view A—A in FIG. 6; and 
     FIG. 8 is a cross-section of the fuel tank wall shown in view B—B of FIG.  6 . 
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     The following description is provided to enable any person skilled in the art to make and use the invention and sets forth the best modes contemplated by the inventor of carrying out his invention. Various modifications, however, will remain readily apparent to those skilled in the art, since the generic principles of the present invention have been defined herein specifically to provide a cryogenic fuel tank having multiple fuel chambers for diverse propellants wherein one set of tank chambers is substantially contained within another set of chambers, and all chambers are arranged substantially conformally within the interior of an aerospace vehicle. 
     The present invention comprises a multi-chamber fuel tank  100  for holding diverse cryogenic propellants. Preferably, the propellant tank  100  possesses a deltoid shape and is designed to be arranged within, and substantially in conformance with, the interior volume of an aerospace vehicle. 
     The propellant tank is a multi-lobed pressure cylinder design that integrates near conformally with the interior volume of an aerospace vehicle. This acts to maximize the propellant volumetric efficiency of the fuselage. The design is comprised of a four lobed pressure cylinder, triangular shaped LH 2  tank that incorporates a two lobed LOX tank partitioned by a common tank wall in the center of the two lobes (see FIGS.  1  and  3 - 5 ). 
     Referring to FIG. 1, it can be seen that the aerospace vehicle  10  possesses a deltoid shaped body portion  12  within which the tank  100  is conformally contained. The tank  100  preferably comprises a plurality of containers or chambers  112 ,  114 ,  116 ,  118  within which it is intended that diverse propellants will be contained. In the embodiment of the invention depicted in FIG. 1, a first pair of the chambers  112  and  118  are arranged parallel to the sides of the body portion of the aerospace vehicle  10 , and are connected together at a forward end portion thereof, but are spaced apart at a rearward portion thereof. The interior chambers  114  and  116  are of a shorter length than the first pair of chambers and are disposed between the first pair of chambers at the rearward end portions of the first pair of chambers. Preferably, a first propellant (such as liquid hydrogen) would be held in the first pair of chambers  112 ,  118 , while a second propellant (such as liquid oxygen) would be contained in the interior chambers  114 ,  116 . 
     FIG. 2 is a side view of the aerospace vehicle  10  depicted in FIG. 1 showing that the chambers of the fuel tank  100  extend substantially the entire length of the vehicle body, and that the interior chambers are shorter in length than the exterior chambers. 
     FIGS. 2A-2D show cross-sections of the vehicle body along the length of the vehicle taken along section lines A—A, B—B, C—C, and D—D, respectively. FIG. 2A is a section taken at a forward region of the vehicle body, and shows that only the exterior chambers  112  and  118  are present. FIG. 2B is a section taken more rearwardly, and shows that the exterior chambers  112 ,  118  are beginning to diverge with facing surfaces  122 ,  128  being separated due to being arranged at an angle to one another. FIG. 2C is a section taken even more rearwardly and shows the two exterior chambers  112 ,  118  and the forward region of the interior chambers  114 ,  116  disposed between the exterior chambers. FIG. 2D is a section taken at the rearward region of the vehicle body, and shows that the four tank chambers are well defined. Preferably, all chambers exhibit at their rearward ends a dome-shaped closure  132  (see the discussion of FIGS. 3-5 below). 
     FIGS. 3-5 show the structural framework  200  used with the tank  100  for the purpose of supporting the chembers of the tank within the interior of the aerospace vehicle  10 . As seen, the framework includes outer panel-like members or septums  202 ,  206  and an inner panel-like member or septum  204 . These septums extend along the length of the aerospace vehicle from front to rear. At the rear of the septums is a curvilinear bulkhead  212 , which extends across the rearward ends of the septums. A bit forwardly of the bulkhead  212  is a first dome-ring member  224  that extends across the septums. The first dome ring member has four circular elements each having a circular opening therethrough for supporting the forward domes of propellant chamber  114 ,  116 . At the forward end of the septums is a second dome-ring  226  which also exhibits circular elements (two) having openings therethrough for receiving and supporting the forward ends of the exterior propellant chambers  112  and  118 . As seen in FIGS. 4 and 5, orthogrid skin panels  232  are welded between the bulkhead  212  and the dome ring member  224 , as well as between the dome ring member  224  and the second dome ring member  226 . Conical dome caps  242  are welded onto the rearward end of the bulkhead  212 , the forward interior end of the dome ring member  224 , and the forward end of the second dome ring member  226  to close the propellant chambers. 
     FIG. 6 is an enlarged view of the cross-section shown in FIG. 2D depicting a first region A—A (shown in greater detail in FIG. 7) where the interior and exterior propellant tanks  112 ,  114  “intersect” and a second region B—B (shown in greater detail in FIG. 8) taken along the “line” of intersection of the interior propellant tank  114  and the exterior propellant tank  112 . 
     The region A—A is shown enlarged in FIG. 7, and it can be seen that a Y-joint structural interconnector  140  has been provided for attaching together the adjacent interior and exterior tanks shown in FIG.  6 . The interior and exterior propellant tanks  112 ,  114  are arranged side-by-side and are attached together at upper and lower points of intersection with the aid of the Y-joint connector  140 . The connector  140  has a central hub  142  and five legs  144  extending from the central hub. Each of the legs includes an enlarged portion or land  146  at the region most removed from the central hub. The enlarged portion of the leg acts as a welding land to facilitate secure attachment of the tank container skin  150 . 
     The region B—B is shown enlarged in FIG. 8 where the preferred construction of the wall or skin  150  of the tank is depicted. The skin  150  comprises a multi-layered structure including a common wall bulkhead  152 , a layer of reusable cryogenic insulation  154  attached the bulkhead, and a thin sheet liner  156  disposed over the layer of insulation. 
     The skin of the container is made of integrally machined orthogrid stiffened panels bump formed to contour. Panel sections are welded together to form each barrel section. Each barrel section is welded to one land of the Y-joint interconnector shown in FIG.  4 . The downstanding leg of the cruciform is used to attach tension septums the full length of the tank. 
     Conical dome sections  132  welded to the aft and forward ends of the chambers incorporate manifolds for propellant feeds, pressurization and fill/drain lines. An internal aft bulkhead is required to react wing bending/torsion loads, main landing gear loads, and engine thrust loads. The orthogrid skin panels provide the necessary stiffness to react vehicle air loads. 
     The tank chambers  112 ,  114 ,  116 ,  118  are constructed from Aluminum 2219 which has the required material properties (high strength and toughness, with no permeability) at cryogenic temperatures. The common tank walls are required to minimize tank surface area (i.e., weight) to achieve the low mass-fraction requirement for an SSTO-RLV. The common walls provide isolation of LH 2  and LOX. To maintain the structural integrity of the tank(s), a positive pressure differential is required at all times (LOX design ullage pressure greater than LH 2  design ullage pressure). Due to the explosive combustion of the propellants, zero defect, zero leak welds are required at the common wall. Friction stir weld technology is a demonstrated engineering solution for leak free welds. 
     Thermal isolation of the two propellants is required to maintain the fuel densities (i.e., to prevent LOX freezing) and minimize different material strains from the propellants at different temperatures (LOX @ −338 degrees F. and LH 2  @ −429 degrees F.). Cryogenic insulation is machined to fill the pockets of the orthogrid panels and then bonded in place. The insulation is secured with a thin sheet panel attached to the orthogrid panel. The common bulkhead tank is pressure stabilized to minimize tank weight. 
     The propellant fuel tank of this invention has been designed to react all vehicle loads, and in so doing, it eliminates the requirement for an intertank structure to transfer inertial LOX loads to a hydrogen tank (common in rocket tank designs). The aft located LOX tank eliminates structural weight associated with a column support of a large inertial mass (i.e., forward LOX design). The tension septums provide hardpoints for the attachment of payload and landing gear, as well as a redundant shear load path for reacting vehicle inertial and air loads. The aft tank domes minimize the duct lengths of the main propellant system. Subsystem routing from the front to the back of the vehicle is accommodated within the three cusps at the tank lobe intersections. 
     Those skilled in the art will appreciate that various adoptions and modifications of the invention as described above can be configured without departing from the scope and spirit of the invention. Therefore, it is to be understood that, within the scope of the appended claims, the invention may be practiced other than as specifically described herein.