Abstract:
Hydrogen powered air vehicles that in some embodiments can fly with very long endurance (10 or more days) at altitudes over 60,000 ft carrying payloads of up to 2,000 pounds. Embodiments may include features such as large wingspan relative to fuselage and an all composite or partial composite structure for light weight and strength. The aircraft of the invention use one or more internal combustion engines adapted for hydrogen combustion, each engine driving propellers. The hydrogen fuel is stored on board in containers, located within the fuselage, as a cryogenic liquid, and is vaporized in a heat exchanger before delivery to the internal combustion engine.

Description:
STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT 
     The United States government has a paid-up license in this invention and the right in limited circumstances to require the patent owner to license others on reasonable terms as provided by the terms of an applicable contract. 
    
    
     TECHNICAL FIELD 
     The present invention relates generally to aerospace technology and more particularly, the present invention relates to hydrogen powered high altitude aircraft. 
     BACKGROUND 
     Certain specialized aircraft are designed to fly long distances and stay aloft for long periods of time. These are sometimes known by the acronym “HALE”—high altitude long endurance. However, the aircraft need not necessarily fly at high altitudes. These aircraft are typically unmanned, remotely controlled and are useful or potentially useful in a variety of applications. For example, in surveillance to conduct aerial patrols along a border, or along a coast of a nation or state, or in remote areas to act as a substitute for a “cell tower” for cellular communications. For such uses, it is desirable that these aircraft have the capability to remain aloft for long periods (more than about three days) and traverse a significant area during that time. Accordingly, the aircraft carry a relatively large quantity of fuel to enable long endurance, whether deployed at high or low altitude. 
     Typically, these aircraft present a large external surface area: they tend to have very large wings relative to fuselage length, typically more than about 3.0. Because fuel is typically stored in the wings, the design of these large wings is generally a compromise between the requirements of “lift” to launch and maintaining the aircraft aloft, and the requirement for fuel storage. In these types of craft there is an additional trade off between reduction in weight (and hence fuel requirements) to extend mission capability versus the increased weight necessary to carry sufficient fuel for long endurance. One approach to meeting these conflicting requirements might be to use solar power but technology to demonstrate flight longer than about 24 hours has not yet been developed. The use of lighter than air stratospheric airships has been proposed but while these achieve high altitude and can stay aloft for days, they are expensive and susceptible to cross winds and so lack the path stability required to carry out missions that require a fixed path of flight. 
     Accordingly, it is desirable to develop an aircraft that could be deployed at high or low altitude that has long endurance and directional stability under most of the expected conditions encountered when aloft. In addition, it would be a benefit if the wings could be designed optimally for flight conditions rather than as a compromise with necessary fuel carrying capability in the wings. Furthermore, other desirable features and characteristics of the present invention will become apparent from the subsequent detailed description and the appended claims, taken in conjunction with the accompanying drawings and the foregoing technical field and background. 
     BRIEF SUMMARY 
     This invention provides unmanned hydrogen powered air vehicles that in some embodiments can fly very long endurance (10 or more days) at altitudes over 60,000 ft carrying significant payloads of up to 2,000 pounds. Embodiments may include features such as large wing spans relative to fuselage (3.0 and greater) and an all composite or partial composite structure for light weight and strength. An embodiment of the aircraft uses one or more internal combustion engines adapted for hydrogen combustion, each engine driving propellers. The hydrogen fuel is stored on board in containers, located within the fuselage, as a cryogenic liquid, and is vaporized in a heat exchanger before charge to the internal combustion engine. In some embodiments, these features enable the craft to provide more than 10 days endurance. 
     Embodiments of the aircraft may include one or more combination or all of several unique features: a hydrogen powered engine; on-board fuel storage of liquid hydrogen in cryogenic containers such as vacuum Dewar fuel tanks; high aspect ratio wings incorporating a high performance laminar flow airfoil; a primarily composite, very light-weight airframe structure; and continuous operation as an unmanned, autonomous flying vehicle for up to about 10 days. 
     Embodiments of the invention can provide unprecedented flight endurance, ranging from over 30 days of endurance capability at altitudes near sea level to 10 days at 65,000 ft altitude, all while carrying a mission communications and payload suit weighing approximately 1,000 lbs. Additionally, embodiments of the air vehicles may include sufficient reliability and/or redundancy of its systems to maintain very-high mission reliability of greater than 90%, if required, over its multi-day flight endurance. The long endurance at high altitude, coupled with its significant payload capacity will allow the air vehicle to provide significantly lower cost solutions in communications, surveillance, data relay and like applications 
     The above and other aspects of the invention may be carried out in one form by an embodiment that is a winged aircraft with a fuselage that encloses a cryogenic container adapted to contain liquid hydrogen; a heat exchanger in fluid communication with an interior of the container, the heat exchanger adapted for receiving liquid hydrogen from the container; and an engine in fluid communication with the heat exchanger, the engine adapted for combustion of gaseous hydrogen as fuel. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       A more complete understanding of the present invention may be derived by referring to the detailed description and claims when considered in conjunction with the following figures, wherein like reference numbers refer to similar elements throughout the figures. 
         FIG. 1  is a schematic perspective view of a high altitude aircraft depicting the long wingspan; 
         FIG. 2  is a partial cross sectional view of an aircraft fuselage according to an embodiment of the invention depicting hydrogen fuel storage containers; 
         FIG. 3  is a block diagram depicting a flow scheme for supplying hydrogen fuel to the engine according to an embodiment of the invention; 
         FIG. 4  is a block diagram depicting a flow scheme of a coolant system for a hydrogen powered aircraft, according to an embodiment of the invention; 
         FIG. 5A  is a cryogenic tank in partial cross section depicting double walled structure and support pillar, in accordance with an embodiment of the invention; and 
         FIG. 5B  is a cross sectional view of a portion of the tank of  FIG. 5A  showing internal insulation. 
     
    
    
     DETAILED DESCRIPTION 
     The following detailed description is merely illustrative in nature and is not intended to limit the invention or the application and uses of the invention. Furthermore, there is no intention to be bound by any expressed or implied theory presented in the preceding technical field, background, brief summary or the following detailed description. 
     The invention may be carried out in one form by an embodiment that is a winged aircraft with a fuselage that encloses a cryogenic container adapted to contain liquid hydrogen; a heat exchanger in fluid communication with an interior of the container and adapted for receiving liquid hydrogen from the container; and an engine adapted for combustion of gaseous hydrogen as fuel, in fluid communication with the heat exchanger. 
       FIG. 1  depicts schematically an aircraft  100  of the type useful in the invention: it has a fuselage  110  and opposed wings  120  extending from the fuselage  110 . The aircraft further has a tail section  125  that includes a rudder  126  and a pair of rear stabilizers  128  (seen more clearly in  FIG. 2 ). The aircraft wings  120  are long in comparison with the length of the fuselage  110 , as is typical in long endurance high altitude aircraft. Typically, but not necessarily, these types of aircraft  100  have a wingspan to fuselage length ratio of 3.0 to about 4.0. While these are characterized as “high altitude” aircraft, meaning they are intended to fly at heights of about 35,000 ft. or more, often 65,000 ft., above sea level, they are also useful at lower altitudes, closer to sea level. The fuselage may comprise a central portion of the aircraft, and various embodiments that include a blended wing body or flying wing do not depart from the scope of the invention. 
     The embodiment shown in  FIG. 1  also has a pair of engines  122 , each driving a propeller  123  and each having a nacelle  124 . Of course, other aircraft embodiments may have a single engine or may have a pair of engines on each wing. Thus, the number of engines is not a limiting factor except that use of a pair of engines provides both redundancy in case of engine failure as well as light weight to conserve fuel for a multi-day long mission. 
     As explained below, the engines  122  are adapted to hydrogen fuel. Fuel is stored on board the aircraft  100  in cryogenic tanks  300 , shown in  FIG. 2  in liquid form, at a pressure of about 30 psi, or other suitable pressure deemed safe and useful. While two tanks  300  are shown, and while these are spherical, one or more tanks of different shapes may be used. As shown, the tanks  300  are contained within a space or cavity formed in the fuselage  110 . In the example, one tank  300  is ahead of the wing attachment location  130 , and the other tank  300  of equal size is rear of wing attachment location  130 . This controls the location of the center of mass of the craft  100  to permit stable flight of the craft. During flight, fuel may be taken from both tanks  300  equally or alternately from each to maintain the center of mass of the craft in a location that ensures craft stability. The nose  150  of the particular aircraft  100  illustrated is adapted to contain control equipment  155  and other equipment ancillary to a mission of the aircraft. 
     In the embodiment of  FIG. 2 , the aircraft  100  is of light weight construction and is made up of a series of skin panels  164  stabilized by a series of underlying supporting frames  160  with longitudinal longerons  162  extending between the frames  160 . Together with the skin  164 , these frames  160  and longerons  162  form the aircraft fuselage  110 . Frames  160  and longerons  162  also provide hard points for the attachment of fuel tanks, landing gear, systems and payload and distribute the loads into the skin  164 . The fuselage shape is selected to conform to a desired (lateral) cross sectional shape of the aircraft that meets several purposes, for example: aerodynamics (drag reduction, etc.) and the mission (size of tanks to carry fuel, for example). Accordingly, the frames  160  are of a shape that are approximately oval in some embodiments, but can vary. To meet the lowest weight criteria while meeting strength requirements, the frames  160  may be fabricated of composite materials, such as carbon fiber impregnated with a suitable engineering polymer such as polyaryletherketone, or epoxy and the like. The frames  160  may alternatively be made of a strong light weight metal alloy, such as an aluminum or titanium alloy. As many longerons  162  as necessary may be used to make a sound structure for fuselage  110 . The longerons  162  may likewise also be of composite materials or may be of a strong light weight metal alloy, such as an aluminum or titanium alloy. Similarly, the skin panels  164  may be of a graphite or fiberglass/epoxy composite sandwich or another composite or light weight metal. Of course, a combination of materials may also be used to take advantage of properties and to achieve a light weight aircraft that meets a reasonable or acceptable cost as balanced against the added fuel consumption and reduced range of a heavier craft. 
     Because embodiments of the aircraft of the invention have hydrogen tanks located within the fuselage, the wings can be designed for optimum performance without any compromises with regard to fuel carrying capacity as was the case with aircraft that have fuel tanks within the wings. Accordingly, embodiments of aircraft in accordance with the invention may have high aspect ratio wings incorporating a high performance laminar flow airfoil that might otherwise not be possible. 
     With regard to the hydrogen-adapted internal combustion engine, any reliable automotive or aircraft engine may be adapted for this purpose by making appropriate modifications. In addition, as with any modern engine, the engine should have an engine control module (“ECU”) that may be modified to take into account the characteristics of hydrogen fuel. Further, to ensure sufficient power, the engine should desirably but not necessarily have compression turbines that increase engine intake air pressure, and these should have intercoolers to remove heat of compression and to permit cooler air feed to the engine, as shown below in the description of  FIG. 4 . 
     As shown in  FIG. 3 , a simplified flow diagram of an embodiment of the fuel system, the fuel tanks  300  each feed into a common outlet conduit  302  that is in communication with suction sides of pumps  310 . The pumps  310  charge liquid hydrogen to a hydrogen reservoir  312  that has appropriate safety features, such as a vent system. Reservoir  312  has a safety vent  311  and is pressured by the pumps  310  to a pressure of about 225 psi, in some embodiments. Liquid hydrogen exiting the reservoir in conduit  313  is monitored for temperature  315  and pressure  316 . The flow in conduit  313  is controllably split into each of two heat exchangers  320  by a pair of inlet control valves  321 . In the heat exchangers  320 , as explained in more detail below, the hydrogen is heated against a warmer coolant to a warmed gas for charging as fuel to the engines  122 . Gas exits the heat exchangers  320  through outlet control valves  322  in conduits  323  and  327  under control of pressure regulators  324  and  328 , respectively, to the engines  122 ,  122 . 
       FIG. 4  depicts an example of a cooling system of the aircraft, in block diagram form, that has three loops: a pair of loops that are identical and that cool the turbine compressors of each engine, and a third loop that maintains the environment for avionics boxes and payload boxes. The cooling system depicted uses a continuously recirculating liquid coolant, such as a mixture of propylene glycol and water, for example. This coolant is continuously circulated to collect heat that must be removed and is then subjected to cooling to remove the collected heat. The coolant is therefore reused in a continuous controlled cycle. 
     In general, heat must be removed from the engines  122 , turbo compression system  171 - 175  of each of the engines  122 , and also from the environmental control system  260 , while heat must be added to the liquid hydrogen to convert it to a gaseous form for combustion. Accordingly, the coolant removes waste heat and applies that heat to the hydrogen gasification. Excess heat from the system is expelled to the environment. 
     Referring to the illustrated example that is based on a two-engine embodiment, three stages of air compression are used for each engine. The coolant system for each engine is identical to that of the other, and so only one will be explained in detail. Ram scoop air enters a low pressure compression turbine  171  and compressed heated air exits and flows to low pressure compression intercooler  172 . The cold coolant is in a cold coolant header  202 . Header  202  has a conduit  203  that diverts cold coolant to low pressure compression intercooler  172 , a heat exchanger which transfers heat from the air to the coolant that exits via conduit  204  to heated coolant header  220 . Cooled compressed air exiting the intercooler  172  enters an intermediate stage compression turbine  173  where it is compressed to a higher pressure and heated as a result of compression. The exiting heated air from turbine  173  enters an intermediate pressure intercooler  174 , where the air is cooled by heat exchange with cold coolant supplied via conduit  205  from header  202 . Warmed coolant exits intercooler  174  as heated coolant in conduit  206  and flows to header  220 . Next, the cooled air exiting intercooler  174  enters a final high compression turbine  175  where it is further compressed and undergoes heating as a result. The heated exiting air from turbine  175  enters a high pressure intercooler  176  that transfers heat from the air to the coolant that enters intercooler  176  from header  202  via conduit  207  and exits the intercooler  176  as heated coolant via conduit  208  to header  220 . Air exiting the intercooler  176  is delivered to an engine intake (not shown) at a pressure and temperature suitable for efficient engine operation. 
     Heated coolant in conduit  220  is drawn to the suction of pump  224  which pumps the hot coolant into a radiator  230  which is air cooled. The coolant is cooled to an extent necessary to retain sufficient heat to vaporize liquid hydrogen and to allow temperature control of the environmental system  260 , illustrated as including the avionics boxes  256  and the payload boxes  258 . Cooled coolant exits the radiator  230  and the flow is controlled to divide into three conduits: one conduit returns to header  202  for use in cooling the turbo compression system (as described above for each engine  122 ) so that the cycle repeats; a second conduit  250  to cool the environmental systems, and a third conduit  270  to supply heat to the liquid hydrogen. 
     As described above, with reference to  FIG. 3 , the liquid hydrogen is generally heated to a gaseous form in heat exchangers  320 . Referring now to  FIG. 4 , the cooled but not cold, coolant in conduit  270  exiting radiator  230  retains sufficient heat to gasify the hydrogen liquid. The coolant enters heat exchangers  320  and exits as further cooled coolant in conduit  272 , having lost heat to the liquid hydrogen. This coolant is routed from conduit  272  into conduit  220 , and hence to pump  224 . Pump  224  sends coolant back to radiator  230 . From radiator  230 , some of the coolant exits to conduit  270  so that the cycle is repeated. 
     A controlled portion of the coolant from radiator  230  entering conduit  250  is charged to a dual core heat exchanger  252 , although other types of heat exchangers may also be used. This radiator-cooled coolant is then circulated by pump  254  via coolant header  255  to the environmental control system  260  that includes avionics boxes  256  and payload boxes  258 , where the coolant picks up heat. The heated coolant circulates back to the heat exchanger  252 , via heated coolant header  259 , where it is cooled. A portion of the coolant exiting heat exchanger  252  enters the suction of pump  254  and is recirculated back to the environment  260 , repeating the cycle. Another portion flows via conduit  253  to conduit  220  for recycle via pump  224  to air cooler  230 . Other arrangements of the cooling system, including systems with a fewer or greater number of components, do not depart from the scope of this invention. 
     As shown in  FIG. 5A , an embodiment of a cryogenic tank  300  depicted in partial cross section has an inner spherical wall  350  and a surrounding outer spherical wall  370  with an annular space  375  between the two walls. The inner wall  350  is supported away from and by the outer wall by a central vertical double-walled pillar  352 . The use of a single support pillar  352  minimizes conductive heat transfer from the outer wall to the inner chamber  355 . The inner chamber  355  contains liquid hydrogen to a height measured by a capacitance gauge  358 , when the tank  300  is in use. The annular space  375  is at a near complete vacuum, with a pressure as low as 1×10 −6  torr, to reduce heat transfer through the annular space. In addition, as shown in  FIG. 5B , in some embodiments the space  375  may be filled with a thermal insulating material, such as for example fiber glass or epoxy, or alternating layers of perforated, double side aluminized polyimide film such as Kapton® 376 (a trademark of Dupont of Delaware, available from MPI Technologies Inc. of Winchester, Mass.) and polyester netting such as Dacron® (a trademark of Invista North America of Wichita, Kans., available from Apex Mills, of Inwood, N.Y.) netting  378 , or other suitable thermal insulation product or combination of products that has a desired low heat transfer. 
     The liquid hydrogen cryogenic tanks  300  should also be fitted with other ancillary and safety equipment as dictated by good engineering practices, for example a vent system to prevent over pressurization, a ground fill and drain system, recirculation system and supply system. 
     In one embodiment, the invention provides an aircraft that has a fuselage and wings extending from the fuselage; a cryogenic container adapted to contain liquid hydrogen; a heat exchanger receiving liquid hydrogen from the container which is in fluid communication with an interior of the container, the heat exchanger; and an engine mounted to the fuselage or wings that is adapted for combustion of gaseous hydrogen as fuel. 
     Optionally this embodiment may include several features such as an intercooled compression system supplying compressed intake air to the engine. The container may be located in a cavity within the fuselage. The container may be a double-walled container with a reduced pressure region approximating a vacuum between the walls. The container may have an inner wall of aluminum alloy. An annular space between the inner wall and a surrounding outer wall may have an insulation material therein. The insulation material may include alternating layers of polyimide and polyester insulating material. The engine may be an internal combustion engine. The heat exchanger may be in fluid communication with coolant that contains heat removed from air compression turbines. The heat exchanger may be configured to transfer heat from the coolant to liquid hydrogen to convert the hydrogen liquid to hydrogen gas. The aircraft may have a pump with a suction in fluid communication with an interior of the container and an outlet in fluid communication with the heat exchanger so that when the aircraft is in use, the pump charges liquid hydrogen to the heat exchanger and the heat exchanger supplies gaseous hydrogen to the engine. 
     In another embodiment, the invention provides an aircraft that includes a fuselage and wings extending from the fuselage; a cryogenic container adapted to contain liquid hydrogen; a pump having a suction in fluid communication with an interior of the cryogenic container and adapted for pumping liquid hydrogen; an engine adapted for combusting gaseous hydrogen as fuel; a cooling system that has a circulating coolant that removes waste heat generated by air compression and hydrogen combustion; and a heat exchanger, in fluid communication with a discharge of the pump and with the coolant, to transfer heat from coolant to liquid hydrogen when the aircraft is in use. 
     Optionally, this embodiment may include several other features including an inter-cooled compression system supplying compressed intake air to the engine. The container may be a double-walled container with a vacuum between the walls. The container may have an inner wall comprised of an aluminum alloy. An annular space between the inner wall and an outer wall may have an insulation material therein. The insulation material may be alternating layers of polyimide and polyester insulation material. 
     In yet another embodiment, the invention provides an aircraft that has a fuselage having wings extending from the fuselage; a cryogenic container adapted to contain liquid hydrogen and located within the fuselage, the container has at least an inner wall surrounded by and spaced from an outer wall, with an annular space between the inner and outer walls having an insulation material, the outer wall supporting the inner wall; a pump having a suction in fluid communication with an interior of the cryogenic container adapted for pumping liquid hydrogen; an internal combustion engine adapted for combusting gaseous hydrogen as fuel; an inter-cooled turbo compression system supplying compressed air to the engine; and a heat exchanger is in fluid communication with a discharge of the pump, and a coolant, to transfer heat from coolant to liquid hydrogen when the aircraft is in use. 
     Optionally, the embodiment may also include a hydrogen reservoir intermediate the pump and the engine, where the reservoir receives liquid hydrogen from the pump and is in fluid communication with the heat exchanger. 
     While at least one example embodiment has been presented in the foregoing detailed description, it should be appreciated that a vast number of variations exist. It should also be appreciated that the example embodiment or embodiments described herein are not intended to limit the scope, applicability, or configuration of the invention in any way. Rather, the foregoing detailed description will provide those skilled in the art with a convenient road map for implementing the described embodiment or embodiments. It should be understood that various changes can be made in the function and arrangement of elements without departing from the scope of the invention as set forth in the appended claims and the legal equivalents thereof.