WO2018138112A1

WO2018138112A1 - A flying craft

Info

Publication number: WO2018138112A1
Application number: PCT/EP2018/051651
Authority: WO
Inventors: Alex Carlton; Richard BINCE
Original assignee: Carpe Astra Limited
Priority date: 2017-01-24
Filing date: 2018-01-24
Publication date: 2018-08-02
Also published as: GB2558949A; GB201701154D0

Abstract

The invention relates to a flying craft (8, 98, 298, 308, 318) comprising lighter than air gas in a receptacle for creating a lifting force by means of buoyancy and a thrust force by combustion in combustion engine as a fuel together with oxygen from an ambient air collector. The invention also relates to a method of operating such a flying craft (8, 98, 298, 308, 318).

Description

A FLYING CRAFT

The invention relates to a flying craft. It is known to couple a rocket to a balloon, to form a so-called "rockoon". For example, the "Bristol Rockoon" consisting of a high altitude balloon coupled to a rocket, in which the balloon lifts the rockoon rocket to around 35 km, allowing the rocket to launch above the majority of the atmosphere of the Earth. This arrangement significantly reduces the aerodynamic loading, and drag forces, on the rocket, which in turn has the effect of reducing the amount of fuel needed to reach the Karman line which commonly represents the boundary between the Earth's atmosphere and outer space at an altitude of 100 kilometres (62 miles) above the Earth's sea level. A reduced fuel requirement means that either a smaller rocket can be used, or a larger payload can be launched.

Another example of a rockoon is the Bloostar by Spanish company Zero2infinity. The Bloostar comprises a helium filled balloon coupled to a rocket. The Bloostar is essentially a launch craft for a nanosatellite. Initially, the helium filled balloon lifts the rocket including a nanosatellite payload to a high-altitude of over 30 km (18.6 miles), i.e. above 99 percent of the Earth's atmosphere. Once in position, the rocket is dropped or decoupled from the balloon and a set of simple liquid-fuelled engines (that are pressure-fed rather than relying on pumps) are fired, launching the rocket including the nanosatellite into orbit. See http://www.gizmag.com/zero2infinity-balloon-rocket- launch/34315/ for more information.

Nissan Motor has a patent, JPS 52118799(A), disclosing a rockoon. Like in the Bloostar, a helium filled balloon lifts the rocket to high altitude, and then the rocket is separated from the balloon and the rocket is launched.

In the case of the Nissan and Bloostar rockoons, the balloon is only partially filled with a small amount of Helium gas at ground level, perhaps roughly only 1% of the maximum volumetric capacity of the balloon. As the finite sized balloon ascends, the gas within the balloon expands to fill the balloon.

A balloon stops rising at a certain height because, for a finite sized balloon receptacle, expansion of the gas within the receptacle of the balloon is limited, which in turn limits the minimal density of the gas within the receptacle. The mass of the balloon receptacle is also a factor. Traditional rockoons like Nissan and Bloostar use the balloon's buoyancy induced Archimedean displacement to reach as high as they can to get above as much atmosphere as possible, so they do not have to deal with the larger amounts of frictional losses on fuel driven motion caused whilst travelling through the lower atmosphere. Launching from height also reduces some gravitational losses. Once the balloon has reached its maximum height, for example at a height of 20 km, the balloon's vertical velocity is zero. The rocket and payload is then decoupled from the balloon, dropped from the balloon in the case of Bloostar. In the case of the Nissan and Bloostar rockoons, the engines are then ignited and the rockets launch with approximately zero velocity from about the 20 KM point. Remco Timmermans reported from the 66^th International Astronautical Congress 2015 in Jerusalem on 15^th October 2015, at the website address

https://twitter.com/search?q=A%20balloon%20to%2020%20km%20MaxQ&src=typd. on the Bloostar presentation, where it had been claimed it is possible to save up to 80% of fuel weight by launching the Bloostar rocket from an altitude of between 20 Km and 40 Km. The exact altitude which the 80% claim relates to is not clear from the comments.

Airships which use a combination of buoyancy and propulsion to lift and fly are also known.

US Patent No. US7614586 (B2) by J.P. Aerospace, discloses an "atmospheric airship". An ATO handout (at http://www.jpaerospace.com/atohandout.pdf) on the J.P. Aerospace website discloses a three-part architecture for using lighter-than-air crafts to reach space, the atmospheric airship being a first stage in this three part architecture. The ATO handout says more about the atmospheric airship. In particular, the atmospheric airship has a combination of buoyancy and aerodynamic lift to fly. The atmospheric airship is driven by propellers. The second part of the architecture is a suborbital space station, a facility parked at 140,000 feet. The suborbital space station is the destination of the atmospheric airship and the departure port for an orbital airship. The orbital airship/dynamic craft uses buoyancy to climb to 200,000 feet. From there it uses hybrid electric/chemical propulsion to slowly accelerate and climb". As it accelerates, it dynamically climbs. In nearly nine hours it achieves orbital velocity. The airship/dynamic craft then flies directly to orbit. The type of gas used to provide buoyancy to the airship does not appear to be disclosed in US Patent No. US7614586 (B2). The static buoyancy of airships in flight is not constant. Changes in weight of fuel on board, due to fuel consumption, have an effect on buoyancy. It can therefore be necessary to control the altitude of an airship by controlling its buoyancy with buoyancy compensation. The Zeppelin LZ 126 airship used Hydrogen, and later helium, as a lifting gas, because of the relative buoyancy of Hydrogen or Helium. Bi-fuel (gasoline and oil) engine driven propellers provided thrust to propel the airship forward. The airship was fitted with equipment to recover water from the exhaust gases from the combustion of the gasoline and oil for use as ballast to compensate for the loss of weight as fuel was consumed, to maintain neutral buoyancy, so avoiding the necessity to vent more expensive Hydrogen or helium.

The LZ 127 Graf Zeppelin airship had engines that ran on "blau gas" or gasoline propellant, and used blau gas as a buoyancy compensating fuel. Blau gas weighs approximately the same as air, its density being only 9% heavier than air. Burning blau gas, and replacing its volume with air from the surrounding atmosphere (because the airship was not airtight), did not significantly lighten the airship, did not significantly change the static buoyancy of the airship, thereby eliminating the need to adjust buoyancy or ballast in flight, as would be required with heavier liquid fuels such as gasoline. The Goodyear Inflatoplane had a fully inflatable fuselage and top-mounted wing. Propeller engines fuelled by a 12: 1 mixture of 80 octane gasoline and SAE 30 oil provided thrust to propel the Inflatoplane.

The V-l flying bomb (also known as the Doodlebug) was an air-breathing rocket (using a pulsejet engine).

The Synergistic Air-Breathing Rocket Engine (SABRE) is a concept under development by Reaction Engines Limited for a hypersonic precooled hybrid air breathing rocket engine. The engine has been designed to propelling the proposed Skylon launch craft, and to achieve single- stage-to-orbit capability. The SABRE engine design comprises a single combined cycle rocket engine with two modes of operation. The air breathing mode combines a turbo -compressor with a lightweight air precooler positioned just behind the inlet cone. At high speeds this precooler cools the hot, ram-compressed air leading to an unusually high pressure ratio within the engine. The compressed air is subsequently fed into the rocket combustion chamber where it is ignited with stored liquid Hydrogen. The high pressure ratio allows the engine to continue to provide high thrust at very high speeds and altitudes. The low temperature of the air permits light alloy construction to be employed which gives a very lightweight engine, essential for reaching orbit. An aim of the present invention is to provide a more efficient flying craft. An aim of the present invention is to provide an improved flying craft and/or at least an alternative, flying craft.

According to a first embodiment there is provided a flying craft comprising a lighter than air combustible gas receptacle, the flying craft being made so that at least at one point in time during its flight the same lighter than air combustible gas provides both buoyancy and thrust to lift the craft, the thrust coming from combustion of lighter than air combustible gas in the lighter than air combustible gas receptacle. Preferably, the flying craft is made so that for a period of time during its flight the same lighter than air combustible gas provides both buoyancy and thrust to lift the craft. Preferably, the flying craft is made so that at a first altitude during its flight and at a second different altitude during its flight the same lighter than air combustible gas provides both buoyancy and thrust to lift the craft.

According to a second embodiment there is provided a flying craft, the flying craft comprising a lighter than air combustible gas receptacle, the flying craft being designed so that, as the altitude of the flying craft increases, a pressure differential between a lighter than air combustible gas inside the receptacle and a gas or gases, such as air, outside of the receptacle, causes a lighter than air combustible gas to be forced out of the receptacle, the flying craft comprising a (rocket) combustion engine or combustion propulsion means, which in turn comprises a combustion chamber, the flying craft being designed to convey a lighter than air combustible gas forced out of the receptacle towards the combustion chamber, the combustion chamber being designed to combust a lighter than air combustible gas from the lighter than air combustible gas receptacle, and to discharge a product or products of the combustion of a lighter than air combustible gas out of an exhaust of the combustion chamber, so as to provide thrust or lift to the flying craft. According to a third embodiment there is provided a flying craft in accordance with Claim 1.

Overpressurization of lighter than air combustible gas (i.e. pressure of lighter than air combustible gas greater than pressure of atmosphere surrounding craft) is the only condition required for lighter than air combustible gas to discharge from receptacle and flow to engine. The term "lighter than air combustible gas receptacle" means a receptacle for lighter than air combustible gas. It will be appreciated the flying craft as defined in the first, second or third embodiment does not specify that the lighter than air combustible gas receptacle comprises a lighter than air combustible gas within the receptacle. In this way, the first and second aspects cover a craft not laden with a lighter than air combustible gas (i.e. a craft before or after it has been used). However, in some embodiments of the invention, the lighter than air combustible gas receptacle of the flying craft as defined in the first, second, or third embodiment comprises a lighter than air combustible gas therein. The term "lighter than air combustible gas" should be understood to cover a combination of different lighter than air combustible gases. Also, the term "lighter than air combustible gas" covers a combination of one or more lighter than air combustible gas(es) and one or more heavier than air combustible gas(es), which combination has an average mass density less than air. In one embodiment, the lighter than air combustible gas is Hydrogen, H₂. In another embodiment, it is Methane, C¾. In another embodiment, it is a combination of Propane with Hydrogen or Methane. It should be noted that Hydrogen provides a passive lift of 8% more than Helium for the same volume used. Moreover, Hydrogen is significantly cheaper than Helium at the time of writing.

The term "lighter than air combustible gas" is used interchangeably with the abbreviation "LTACG" throughout the specification. The lighter than air compbustible gas receptacle is preferably rigid.

The term "passively provide buoyancy" refers to the inherent buoyancy of the LTACG, prior to, or without, using the LTACG in combustion.

The flying craft is designed to transfer a fuel from the receptacle to the (rocket) combustion engine or the combustion propulsion means as a direct consequence of an increase in altitude of the flying craft. Due to the operational differences between a craft in accordance with the invention, and prior art crafts, it is estimated that a craft in accordance with the invention will have an improved fuel efficiency. If the lighter than air combustible gas is Hydrogen, the flying craft in accordance with the invention is still travelling at a significant velocity when it gets to for example 20 KM, because Hydrogen has a lower density than Helium so provides more buoyancy per volume of gas. The flying craft will be carrying sufficient LTACG to be used as a fuel that at the 20 KM point the flying craft in accordance with the invention gets the same advantages as the Nissan and Bloostar systems, plus a significant velocity. Also, the flying craft in accordance with the invention utilises a lighter than air combustible gas instead of non combustible helium gas used by Nissan and Bloostar. The flying craft in accordance with the invention utilises the changing environment in the atmosphere as it ascends to displace LTACG into the engine. Complex pumps are not needed to move fuel to the engine, although certain more advanced models of the craft could use simplified pumps for greater control of ascent profiles. Whilst the LTACG buoyancy contributes towards lifting the craft, when some of the LTACG is used as a fuel more of its energy goes into acceleration of the craft because the craft does not need to waste as much energy overcoming gravity. A minor extra efficiency comes from the downward force of the LTACG being expelled from the balloon part into the engine due to the existing pressure differentials causing a small non negligible amount of extra thrust. The helium gas used by Nissan and Bloostar is also more expensive than a lighter than air combustible gas.

In one embodiment of the invention, the flying craft consists of a rockoon. In another embodiment of the invention, the flying craft comprises a rockoon. The flying craft in accordance with the invention is more than a balloon coupled to a rocket, or a balloon connected to a rocket, the flying craft in accordance with the invention is a truer "hybrid" of a balloon and a rocket, where energy from lighter than air combustible gas inside the balloon receptacle is transferred to the rocket, and becomes fuel for the rocket, in a synergistic manner.

The flying craft travelling at a significant velocity when it gets to for example 20 KM allows the craft to reach a significant proportion of orbital/escape velocity for space related purposes, or to reach a significant proportion of suborbital velocity for terrestrial travel purposes. The craft preferably comprises an inert gas receptacle which surrounds at least part, the majority, or substantially all of the LTACG receptacle, preferably the receptacle may have an inert gas therein, most preferably the inert gas is Helium. Such a craft arrangement reduces safety concerns associated with the combustible nature of a LTACG, especially Hydrogen. The engine may be actively cooled to remove excess engine heat, consequently the engine can be smaller and lighter, thereby increasing the efficiency of the craft.

The flying craft is preferably made so that, at an altitude of zero, i.e. at ground level, a pressure differential between a lighter than air combustible gas inside the a lighter than air combustible gas receptacle and gas or gases surrounding the outside of the craft, can exist, so as to cause a lighter than air combustible gas to be forced out of the receptacle, via the reaction chamber, towards the combustion chamber of the combustion engine or combustion propulsion means. The craft receptacle is preferably fully inflated with LTACG at ground level. Hence, the craft can lift significantly more than a traditional rockoon. It will be appreciated that if the craft in accordance with the invention waited until too high an altitude before permitting LTACG to leave the receptacle (towards the engine or propulsion means), the relatively higher pressure of the LTACG compared to the surrounding low pressure atmospheric environment could cause the receptacle to be damaged by it bursting. Accordingly, the craft engine or propulsion means preferably is designed to combust LTACG from the receptacle at any altitude from takeoff onwards, thereby giving the craft the capability to prevent dangerously high pressure build up in the LTACG relative to the atmospheric surroundings. Obviously the LTACG could be vented to atmosphere to avoid high pressure build up but that would be wasteful. The receptacle and the (rocket) combustion engine or the combustion propulsion means are designed to be in gaseous communication. However, it is envisaged some embodiments of the flying craft can comprise a valve or regulating means to control the flow of fuel from the receptacle to the rocket combustion engine or the combustion propulsion means. In other words, transient closing of the flow of fuel from the receptacle to the (rocket) combustion engine or the combustion propulsion means is envisaged in some embodiments of the flying craft.

The term "an Oxygen receptacle" means a receptacle for Oxygen. The term "Oxygen collector" is intended to cover the flying craft comprising a physical structure intended to permit the collection of a gas such as air, for the extraction of, or creation of, Oxygen from the collected gas. If an Oxygen receptacle is provided, the flying craft may comprise Oxygen, or in other words be provided with Oxygen.

A product or products of the combustion of fuel and Oxygen is/are (mostly) air and water. The craft may be heavier than air at ground level. The craft is preferably lighter than air at ground level. The craft may become heavier than air during flight.

If a liquid ballast receptacle is provided, the flying craft may comprise liquid ballast, or in other words be provided with liquid ballast, such as, but not limited to, water.

In one embodiment of the invention, the flying craft comprises an aircraft. In other embodiments of the invention, the flying craft consists of, or comprises, an airship. In one embodiment of the invention, the flying craft comprises an air breathing engine. In that way, the craft need not carry oxidiser, and efficiency gains can be had.

According to a fourth embodiment there is provided a method of operating a flying craft in accordance with Claim 28. One or more optional method steps corresponding to features disclosed in the claims relating to the first, second, or third embodiment, and/or disclosed in the description are envisaged.

Other optional and preferred features of the invention are set out in the dependent claims, and the description, below. The features of one embodiment can be combined in any complimentary manner, with one or more features of embodiment, the dependent claims, and/or with one or more features of the description, where such a combination of features would provide a working embodiment of the invention.

Various flying crafts in accordance with the invention will now be described, by way of example only, with reference to the accompanying drawings, in which:

Figure 1 is a schematic view, in about 120 degrees cross section, taken in the direction of arrows I-I in Figure 2, of a first flying craft in accordance with the invention, termed the Potent Voyager Craft by the Applicant. The "about 120 degrees" angle is actually less than 120 degrees by half the circumferential extent of one of the air scoops about the circumference of the craft, whatever size the sir scoops are. The air scoops will be described hereunder.

Figure 2 is a schematic plan view, in cross section, in the vicinity of an engine of a Potent Voyager

Craft of Figure 1, taken along line II - II in Figure 1 , viewing in the direction of the arrows. Figure 3 is a schematic plan view, in cross section, in the vicinity of an engine of a Potent Voyager Craft of Figure 1, taken along line III - III in Figure 1 , viewing in the direction of the arrows.

Figure 4 is a schematic side view, in cross section, exploded in the vicinity of the engine of the Potent Voyager Craft of Figure 1 to show more detail.,

Figure 5 is a schematic plan view from below, in cross section, in the vicinity of the engine of a Potent Voyager Craft of Figure 1 , taken along line V - V in Figure 1 , viewing in the direction of the arrows.

Figure 6 is a schematic side view, in cross section, of another flying craft in accordance with the invention, termed the Coanda Craft by the Applicant.

Figure 7 is a schematic plan view from above, in cross section, of the Coanda Craft of Figure 6 taken along line VII - VII in Figure 6.

Figure 8 is a schematic view, from above, in cross section, taken in the direction of arrows VIII- VIII in Figure 6, exploded in the vicinity of the engine of the Coanda Craft to show more detail of the central part of the craft located radially within lines VIII- VIII in Figure 7.

Figure 9 is a schematic view, in 180 degrees cross section, taken in the direction of a line IX-IX in Figure 8 and Figure 10, taken from the bottom of Figure 6 of the drawings looking upwards, exploded in the vicinity of the engine to show more detail.

Figure 10 is a schematic plan view, from below, in cross section, in the vicinity of the engine of the Coanda Craft of Figure 6.

Figure 11 is a schematic side view, in cross section, of a safety valve assembly.

Figure 12 is a schematic side view, in cross section, of an alternative safety valve assembly.

Figure 13 is a schematic view, from above, in cross section, of the part of the Coanda craft of Figure 8, zoomed in relative to Figure 8, and showing a first outer wall coolant flow system, Figure 14 is a schematic view, from above, in cross section, of the part of the Coanda craft of Figure

8, zoomed in relative to Figure 8, and showing a second outer wall coolant flow system,

Figure 15 is a schematic view, from above, in cross section, of the part of the Coanda craft of Figure 14, zoomed in relative to Figure 14, and showing an inner wall coolant flow system,

Figure 16 is a schematic view, in 120 degrees cross section, in some respects like the view of Figure

9, which shows the first outer wall coolant flow system, the second outer wall coolant flow system, and the inner wall coolant flow system,

Figure 17 is a schematic view from the front, in cross section, of another flying craft in accordance with the invention, termed the Liberator Craft (Potent Voyager Craft class) by the Applicant,

Figure 18 is a schematic view from the front, in cross section, of another flying craft in accordance with the invention, termed the Liberator Craft (Coanda Craft class) by the Applicant,

Figure 19 is a schematic view from the front, in cross section, of another flying craft in accordance with the invention, termed the Liberator Craft (combined Potent Voyager Craft and Coanda Craft class) by the Applicant,

Figure 20 is a perspective view from one side of the Liberator Craft (Potent Voyager Craft class) of Figure 17, and

Figure 21 is a perspective view from one side of the Liberator Craft (Coanda Craft class) of Figure 18.

The applicant has devised a range of efficient launch vehicles that use lighter than air combustible gases such as Hydrogen and/or Methane. Gaseous Hydrogen H₂ 143 MJ/kg is the preferred single gas option. Gaseous Methane 55.6 MJ/kg is an alternative. Gaseous Hydrogen and gaseous Methane have differing advantages and disadvantages but each is still lighter than air. A combination of Gaseous Hydrogen and gaseous Methane could also be used, with a mixture of the advantages of both. All calculations below are based on pure Hydrogen. The use of lighter than air combustible gas(es) enables the gas(es) to be used for both a passive lift and active lift fuel system in atmosphere. The passive lift comes from the positive buoyancy of lighter than air gas(es) whilst the launch vehicle is in the atmosphere (utilising Archimedes displacement principle). The engines of the crafts are also "air breathing" in the atmosphere.

While the entire launch vehicle (along with payload) is lighter than the air mass displaced, it will be getting greater than 9.81 Newtons per Kg of launch vehicle countering gravitational drag with anything over 9.81 Newtons per Kg accelerating the vehicle on a per second squared basis, until it reaches so called "terminal buoyant velocity" (t.b.v.). Terminal buoyant velocity is the terminal velocity for a passively buoyant body that rises within the pressure differentiated medium it is in. Unlike terminal velocity that falls in the pressure differentiated medium that it is in, for example :-

• Terminal velocity decreases as the body falls deeper into the medium (atmosphere) as the pressure increases.

• Terminal buoyant velocity increases as the body rises in the medium (atmosphere) as the pressure decreases. (This is independent of the decreasing buoyancy of the body as the body is rising within the medium). When it reaches the height of equilibrium (when launch vehicle along with its fuel at this stage and payload mass equal air mass displaced), gravitational drag of 9.81 newtons per Kg is directly countered. The height of equilibrium is an optimisable parameter dependent upon the flight profile, by changing the rate of usage of the various fuels. Above the height of equilibrium, gravitational drag is only partly compensated in a diminishing curve while the vehicle is still in the atmosphere.

Active lift in the vehicle comes from:- 1) The explosive combustion of pressure differentially displaced lighter than air combustible gases such as Hydrogen and or Methane from a fixed volume balloon into an air breathing engine.

2) Water to be used as ballast and as an engine coolant to reduce engine mass to save weight and to give thrust from wasted heat, in the form of super heated steam. The steam provides thrust in accordance with Newtonian laws, and the steam is vectored accordingly to the thrust direction requirement. This increases the overall efficiency of the system, and gives extra energy density. The specific heat capacity of H₂0 is 4.18 joules per gram per 1 degree centigrade, therefore 4.18 Mega joules per metric ton of H₂0 per 1 degree centigrade, therefore 418 Mega joules are required to produce ordinary steam for a metric ton of H₂0, therefore 4.18 Giga joules are required to produce superheated steam at a 1000 degree centigrade for a metric ton of H₂0. This means that only a relatively small amount of H₂0 is required to cool the engine and give extra thrust in the form of superheated steam. The energy used is that of the waste heat from the primary thrust reaction. This increases the overall energy efficiency of the system.

3) An optional extra, Hydrogen peroxide (HTP(H₂0₂)) to be used as ballast, engine coolant and oxidiser which gives thrust in the form of super heated steam, produces Oxygen that then explosively combusts with the pressure differentially displaced lighter than air combustible gases that will produce extra thrust. This also gives extra energy density.

4) There is some additional active lift from the pressure differentiated displaced lighter than air gases. For example, this is produced in a similar manner that a child's party balloon rocket works. For instance some of the thrust is obtained just from the expulsion of the gases. This will probably be only less than 1% of the overall lift. Therefore this is a small but significant additional thrust that needs to be accounted for in the calculations of the trajectory of the vehicle.

5) Numbers 2 and 3 are the preferred environmentally clean ballasts, for extra energy density and engine coolants but a range of other fuels could be used in their place, in this hybrid part of the engines. In a similar way that diesel engines can run on a range of differing fuels.

It has been calculated the minimum energy to get lKg mass of payload to space with no atmospheric or gravitational drag is 17.8 KwHr or 64.08 mega joules for escape velocity. This was calculated for "space elevator" use. As lKg mass of H₂ burnt in air has a theoretical maximum of a 143 mega joules per kilogram. Therefore lKg of gaseous Hydrogen mass has a theoretical capability of accelerating itself and lKg of launch vehicle/payload to escape velocity with a margin of error of 14.84 mega joules theoretically to take account of the reduced atmospheric and gravitational drag due to the ascent profile of the crafts in accordance with the invention. The margin of error of 14.84 mega joules can be calculated in one of three different ways set out by Jim Cline full at http://home.earthlink.net/~jedcline/ets.html. The first calculation method is based on the kinetic energy added to the launched mass; the second way is to use a simplification described by Arthur C. Clarke decades ago, and the third way is to use the gravitational equation. The first calculation method, the kinetic energy method, arguably the simplest version of the calculation, follows below.

The convention of using the symbol "E" followed by a number, to designate the number of zeroes to move the decimal point to the right, in a scientifically notated number, will be used here. For example, " 1.0E6" would be 1,000,000; and " 1.0E-2" would be 0.01.

Under the kinetic energy calculation, the energy required to accelerate one pound mass to the 25,000 mph escape velocity is: It is easier to first use the metric system for such calculations, for 1 Kg accelerated to escape velocity as follows:

25,000 mph = (2.5E4 mph)(0.447 m/s per mph) = 1.1175E4 m / s. W = ((1 Kg)(1.1175E4)2) / 2 = 6.24E7 Joules.

Since 1 Joule = 2.78E-7 KwHr, the work needed to accelerate 1 Kg to 25,000 mph is 17.8 KwHr.

Since 1 Kg = 2.2 lbm, then it would theoretically take 7.89 KwHr to accelerate one pound mass to 25,000 mph, escape velocity.

The 17.8 KwHr can then be put through a conversion program many of which are available on the internet to find that it is equivalent to 64.08 Mega Joules. So lKg of Hydrogen, which is known to produce 143 Mega Joules when burnt in air, could lift a Kg of itself and a Kg of payload with 14.84 Mega Joules to spare as follows:

143.(64.08+64.08)

It is likely the 143MJ refers to liquid Hydrogen. By using ambient temperature Hydrogen in its gaseous form more energy will be available. Therefore, the spare energy in the above figures is likely to be conservative.

Part of the efficiency of the applicant's new craft, form of transport to get into Low Earth Orbit (L.E.O.) and above, is gained by displacement, i.e. it is intended to make the whole vehicle lighter than air at ground level by using a fixed volume Hydrogen structurally gas impermeable balloon(s), with optional helium sheath protective and structural strong gas impermeable balloon(s), using the Hydrogen in its gaseous form as fuel on the way up. The structural integrity of the craft will be maintained by the helium sheath to prevent any concertina effect. The pressure of the helium will be at least equal to or greater than the combustible lighter than air gases, to prevent any pressure differential osmotic flow into the helium sheath, and to give greater structural strength.

With traditional rocketry, a large amount of effort goes into lifting the weight of the vehicle it's payload and fuel against gravity over the time of the flight i.e. you need to expend 9.81 Newtons of force for every kilo of mass of your whole launch vehicle, payload and fuel for every second of flight until you reach orbital velocity.

When the applicant's new craft is lighter than air, all thrust from the engine goes into accelerating the vehicle, rather than the 9.81 newtons per kilo per second that just go into holding it in place against gravity with normal rocketry.

The primary fuel in the applicant's new craft (when it is switched off, whilst it is not expending primary fuel) passively lifts the whole vehicle with a force of just over 12 Newtons for every cubic metre of gaseous Hydrogen at STP (Standard Temperature and Pressure - defined as air at OoC (273.15 K, 32oF) and 1 atm (101.325 kN/m2, 101.325 kPa, 14.7 psia, 0 psig, 30 in Hg, 760 torr). By way of reference, NTP - Normal Temperature and Pressure - defined as air at 20oC (293.15 K, 68oF) and 1 atm ( 101.325 kN/m2, 101.325 kPa, 14.7 psia, 0 psig, 30 in Hg, 760 torr).

There will be a height at which the Hydrogen will stop lifting the whole vehicle's weight against gravitational losses, but this is a slowing diminishing amount of lift rather than a binary cut off.

Rockets traditionally carry their own oxidiser often LOX (Liquid Oxygen) which can be dangerous to handle as well as a weight penalty along with all its ancillary equipment. You would need eight times the mass of Oxygen to burn one mass of Hydrogen at 100% efficiency using liquid Hydrogen as a fuel, if it could be done. As far as it is known, rocket engines cannot burn Liquid Hydrogen and LOX at 100% efficiency.

One of the unique parts of the Engines of applicant's new craft is that both the primary fuel and oxidiser are starting in their gaseous states. Thereby this gives the opportunity to achieve a 100% burn efficiency, unlike solid or liquid reactions, due to using the nitrogen in air as a chaperone molecule for the reaction.

Energy density

The flying craft is made so that the pressure of the LTACG is either a) equal to the atmospheric pressure surrounding the craft at any given altitude, or b) greater than the atmospheric pressure surrounding the craft at any given altitude. At launch, the LTACG is at 1 atmospheric pressure, i.e. 1 bar. Over pressurization of the LTACG, i.e. a LTACG pressure greater than the atmospheric pressure surrounding the craft at any given altitude gives more flexibility of flight profile. However, the pressure can be equalized inside and outside the craft at times whilst the craft is ascending. For the specific embodiment of the craft disclosed over pressurization of the LTACG should not exceed 14 bar, as explained below. Energy density for the flying crafts in accordance with the invention (also referred to as the RAROSS system herein) can be increased for reactive lighter than air gases, (e.g. Hydrogen and/or Methane) by increasing pressure. For Hydrogen the increase in energy density, that is the amount of fuel can be up to 14 bar when the outside environment is at STP which would still maintain some passive lift close to neutral buoyancy.

For example, the table below shows the varying mega joules of energy and passive lift in kilograms available, from 1 to 14 bar in a fixed volume balloon sized to initially contain IKg mass of Hydrogen at STP. 1 Kg H2

143MJ 143

1 1.12595 cubic Meters 1 1.12595

H2 is 0.08988 Kg per cubic meter 0.08988

Passive Lift From a cubic meter of H2 1.202 Kg 1.202

Cubic meter of air 1 .292 Kg 1.292

Volume of Balloon 1 1.12595

No. Bar Kg mass of H2 Mega Joules Passive lift Kg

1 1.000000386 143.000055198 -13.374727014

2 2.000000772 286.000110396 -12.374726628

3 3.000001158 429.000165594 -1 1.374726242

4 4.000001544 572.000220792 -10.374725856

5 5.00000193 715.00027599 -9.37472547

6 6.000002316 858.000331188 -8.374725084

7 7.000002702 1001.000386386 -7.374724698

8 8.000003088 1 144.000441584 -6.374724312

9 9.000003474 1287.000496782 -5.374723926

10 10.00000386 1430.00055198 -4.37472354

11 1 1.000004246 1573.000607178 -3.374723154

12 12.000004632 1716.000662376 -2.374722768

13 13.000005018 1859.000717574 -1.374722382

14 14.000005404 2002.000772772 -0.374721996

In the table above, the Applicant selected a LTACG receptacle volume of about 11 cubic metres because such a volume would contain IKg mass of Hydrogen at one atmospheric pressure. IKg mass of Hydrogen is the base unit for our energy calculations. When IKg mass of Hydrogen is burnt efficiently it would provide 143 mega joules of usable energy. This would at 1 atmospheric pressure be able to lift 13.374 Kg of weight passively. The same volume receptacle could contain 14 Kg mass of Hydrogen providing 2 giga joules of usable energy, but that would only give 0.374 Kg weight of passive lift. It is not envisaged to have a craft LTACG receptacle volume as small as of about 11 cubic metres, but as the craft LTACG receptacle volume (and therefore the craft) increases in size the weight of the crafts balloon component only increases with its surface area. However the energy density and passive lift increases with volume. At larger scales of craft size the 14 bar 0.374 Kg weight passive lift per of about 11 cubic metres would be able to lift the craft as its weight would be proportionally less. Therefore if the craft is large enough the passive lift can result in actual lift for said craft even at 14 bar. In some circumstances the passive lift may provide only a contribution towards overall active lift, the thrust of the engine providing the remainder of the lift. The table is proof that passive lift is available for Hydrogen, from 1 to 14 bar, as long as the containing fixed volume balloon's weight is less than the passive lift available. It will be appreciated from the table above that LTACG pressure above 14 bar would result in a negatively buoyant craft. The reason why this is of importance, is that in traditional rocketry it is paramount to achieve maximum energy density to reach the maximum amount Isp (i.e. fuel efficiency, discussed below) possible. This is because of the weight of the launch vehicle and its fuel, act as penalty against the energy used to achieve lift. The disadvantage of a fixed volume balloon (containing the lighter than air gaseous fuel) is that the medium (atmosphere) pressure levels drop off with height, causing the amount of buoyancy to drop off as the vehicle ascends. This can be partially compensated by over pressurisation.

For instance, the dumping of over pressurised lighter than air combustible gases (Hydrogen and/or Methane), through and out of the vehicle, actually lightens the vehicle. This is the reason why both the primary and secondary fuel/coolant acts in the same way that ballast operates in traditional ballooning, by getting rid of excess weight, it increases the buoyancy at the level of dumping.

The advantage of over pressurising is that the optimal buoyancy level can be maintained in the different pressure levels of the medium (atmosphere) that the vehicle is ascending in, for a longer period by choosing the rate of expenditure of the over-pressurised gases, giving a greater overall efficiency by maintaining buoyancy and by avoiding excessively exceeding terminal buoyant velocity causing losses. Despite the logarithmic drop off of buoyancy with height in the ascent profile, i.e. For instance the same volume of the balloon fuel cell (containing the lighter than air gaseous fuel) would only displace approximately half the mass of air mass displaced at ground level at the height 5.5 kilometres. Therefore excess pressurisation and active ballasts would be turned into velocity, momentum and acceleration of the vehicle, whilst maintaining optimal buoyancy for a range of heights for the flight profile.

Another advantage of over-pressurising of the buoyant gases, is that it can be used as a way to provide supersonic gas injection into the combustion chamber. This could considerably reduce backblast losses and increase the final exit velocity of the combusted gasses increasing overall efficiency.

So to conclude, in traditional rocketry any extra mass is a weight penalty, possibly decreasing overall energy density, but using lighter than air reactive gases increases usable energy availability as there are more and more of the gases, providing extra lift both actively and passively.

The pressure differential between the LTACG and the outside atmosphere can be controlled at any point in time by using a valve or valves. Separately the pressure differential at any instance in time can also be controlled positively, negatively or to maintain equilibrium by changing the rate of use or not of ballast, secondary fuel and temperature control.

The LTACG negates craft weight and provides thrust, and the helium negates craft weight, but the negation of weight in all cases also increases the efficiency of any form of fuel used.

The crafts in accordance with the invention can use an air breathing engine. Using air, has the advantage of it also containing nitrogen acting as a chaperone molecule for the chemical reaction between Hydrogen and Oxygen. With a lean burn fuel mix, it is expected to get 100% burn from the fuel, with the exhaust being mostly air and water in the form of super heated steam, which gives the vehicle thrust.

Rocket fuel efficiency is measured in Specific impulse, usually abbreviated I_Sp. The Challenger space shuttle had a more or less constant I_SP of about 453 seconds over a range of Mach speeds 0-10 on the propulsion performance diagram (see diagram on page 78). An air breathing gaseous Hydrogen engine has a theoretical maximum Isp of 8,000 diminishing to about 6,000 from Mach 6 - 10 and, below Mach 6, it is off the chart above it on the same diagram. With the best design and constructed version of the crafts in accordance with the invention it can be possible to hit a high percentage of the theoretical maximum. Due to the way Isp is calculated whilst the vehicle/payload is lighter than air the maximum theoretical Isp increases considerable above the 6-8000 range, because the lighter than air gases lift and accelerate themselves and the vehicle/payload passively, with nearly all the active acceleration going into just the vehicle/payload and not needing to lift the lighter than air gaseous fuel. See formal proof on page 72. The primary three types of the applicant's new craft are 1. The Potent Voyager, 2. The Coanda, and 3. The Liberator.

All three classes of vehicles will be streamlined for as low a Reynolds number as possible. Although frictional losses are partially countered by the increasing reactive efficiency caused by the heating of the fuel from the friction incurred in these cases. By using a helium sheath the effect of the thermal transfer properties are protective of the outer balloon envelope, over the thermal heating degradation of said envelope, and also transfer the heat efficiently and safely into the inner lighter than air combustible gaseous fuel balloon envelope. Thereby increasing the efficiency of the lighter than air combustible gaseous fuel reactions.

Figure 1 shows a craft 8, called The Potent Voyager class by the Applicant. The craft 8 is for the launch of small and micro satellite payloads. The Potent Voyager class craft 8 is shaped like a traditional rocket. However, instead of, from top to bottom, payload, fuel and engine, the configuration of the craft 8 will be, from top to bottom, fuel, payload, (protective active ballast) and engine. This configuration will ensure safety and inherent stability, because the lighter than air gas in the fuel section passively holds the system stably upright. In contrast, traditional rockets often have toppling problems in the early stages after lift-off.

The craft 8 comprises the following parts:

12. An outer envelope of an aerodynamic shaped gas impermeable balloon, acting as a temperature management system, and as structural support for the receptacle or tank 18. The outer envelope 12 can be constructed using aluminium-ised Biaxially-oriented polyethylene terephthalate (BoPET) more commonly known as Mylar™, Melinex™ and Hostaphan™, which is set to reflect heat outwards away from the craft. The outer envelope 12 may be laminated, with other fabrics using similar techniques that are used to make space suit fabrics, for protection against rips, punctures and tears, for structural strength and safety according to the flight profile required. The outer envelope 12 can be constructed, for example, by a lamination of, from outer layer to inner layer: - Polyethylene terephthalate metallised (with titanium) film, silica gel adhesive, Kevlar™ like material, silica gel adhesive and a layer of Polyethylene terephthalate metallised (with aluminium) film. This should provide a strong light weight envelope that should be resistant to punctures and self sealing for small punctures, if they should occur. The process required has to take into consideration the thermal transfer properties of the materials used.

14. An inner envelope of a gas impermeable balloon, acting as a temperature management system, and used as structural support, and also as a fuel container, receptacle or tank for a primary lighter than air combustible gas 18 and/or gases 18. The inner envelope 14 can be constructed in a similar manner to as that of the outer envelope 12, but with an emphasis on thermal temperature transfer without the need for as much structural strength. Also, the aluminium-ised BoPET is set to reflect heat inwards.

Other materials may also be considered, for production models. For example, in due course, if and when scaled up production of graphene/ advanced nano diamond is readily and economically available, the outer envelope 12 and the inner envelope 14 can be constructed of a graphene/ advanced nano diamond (on a plastics substrate).

18. A primary lighter than air gaseous fuel, e.g. H₂ (Hydrogen) and/or CH₄ (Methane). To be used to passively and actively lift the craft 8.

10. A reinforcement for re-enforcing of the outer envelope temperature management system acting as additional heat shielding on the leading edge of the outer envelope temperature management system 12. The reinforcement 10 can be constructed in a similar manner to supersonic inflatable aerodynamic decelerator (SIAD), which has been tested successfully by NASA, at supersonic velocities. However, where SIAD uses nitrogen, in the invention, it is preferable helium is used instead, because it is more buoyant and less reactive. Also, in the invention, it is preferable that a more aerodynamic profile is used. See NASA websites for more information. It will be appreciated that SIAD is just one of several possible options available.

16. Helium gas, which will be used to inflate the gas impermeable balloon formed by outer envelope 12 and inner envelope 14 at a pressure to provide structural strength, when needed. The Helium gas 16 forms part of the temperature management system and can transfer frictional heat energy, due to drag on the craft 8, to a lighter than air gaseous fuel 18 in a controlled and safe manner. The outer envelope 12, inner envelope 14, the reinforcement 10 on the leading edge of the outer envelope 12 and the Helium 16 between the outer envelope 12 and inner envelope 14 may be optional if appropriate pressures are used, and if a suitable advanced material can combine strength, gas impermeability and heat energy transfer in a light weight structural layer in order to house a primary lighter than air gaseous fuel 18 such as Hydrogen, and function as an outer shell of the craft 8.

20. A fuel line for primary lighter than air gaseous fuel 18 displaced by pressure differentials from the fuel container, receptacle or tank 14 towards an engine 36. The fuel line 20 can be constructed from High-density polyethylene (H.D.P.E. ).

21. An outer hull of a structure containing the payload and ballast/secondary fuel tanks (which can be made from H.D.P.E.), can be constructed from carbon fibre, and can be glued and/or welded to the outer envelope 12 of the gas impermeable balloon.

22. Payload.

24. A safety valve. The safety valve is upstream of outlet of main fuel line 20. Examples of a safety valve are described with reference to Figures 11 and 12. The safety valve can be an integral part of a 3D printed alumina engine 36. Alternatively/additionally, a traditional type of safety valve can be employed. The safety valve 24 may be optional depending on the flight profile required.

26. Receptacle for a ballast/secondary fuel/LTACG coolant. The ballast/secondary fuel/LTACG coolant can be rocket grade Hydrogen peroxide known as H.T.P., which has an approximate chemical formula H₂0₂. The terms HTP and H₂0₂ are used interchangeably throughout. The ballast/secondary fuel/LTACG coolant can be water (H₂0). H₂0 has more cooling and less extra thrust. H₂0₂ has more extra thrust and less cooling. Separate receptacles can each house H₂0 and H₂0₂. The decision to use H₂0 or H₂0₂, or both, would depend upon the flight profile required. 28. One or more Hydrogen and/or Methane branch pipe. In the embodiment of the invention shown, there are three Hydrogen and/or Methane branch pipes as shown in Figure 3. The pipes are spaced at 120 degrees intervals about a longitudinal axis of the craft, The pipes can be an integral part of a 3D printed alumina engine. 30. One or more downpipe for H₂0₂ and/or H₂0 from receptacle 26. In the embodiment of the invention shown, there are three downpipes as shown in Figure 3., the downpipes are spaced at 120 degrees intervals about a longitudinal axis of the craft. The downpipes can be an integral part of a 3D printed alumina engine.

32. One or more air scoops for combined atmospheric air intake and compression. In the embodiment of the invention shown, there are three air scoops, spaced about the circumference of the outer envelope 12 at 120 degrees intervals, as shown in Figures 2 and 3 and 5. The air scoops can be an integral part of a 3D printed alumina engine, possibly reinforced with carbon fibre. The air scoop 32 outer wall inclines at an angle of between 60-70 degrees to a horizontal line across the sheet.

34. Receptacle for H₂0 ballast for engine coolant and additional thrust from the steam produced by its use as a coolant preventing the engine from overheating, which assist in the active lift of the craft. The receptacle 34 has a conical shape. Steam outlets omitted from Figure 1 for conciseness, but steam outlets are shown in Figures 4 and 5. The receptacle 26 and the receptacle 34 are interchangeable features that can also be used in combination. The receptacle 26 and the receptacle 34 may be optional depending on the flight profile required. 36. A Bell engine. Expanded details shown in Figures 3, 4 and 5. The engine can be a 3D printed alumina engine.

38. Bell engine exhaust chamber. The exhaust chamber, possibly reinforced with carbon fibre, can be an integral part of a 3D printed alumina engine 36.

Note: Not all outlets and inlets shown. For example, steam exhaust holes (described below under reference numerals 64 and 72) are omitted from Figure 1 due to the scale of the diagram of Figure 1, but steam exhaust holes 64 and 72 are described in connection with Figures 4 and 5. Figure 2 shows:

40. A top of the fuel line 20 (see Figure

The Bell engine exhaust 38 is not shown in Figure 2. Figure 3 shows:

42. An outer wall of engine 36. The outer wall can be an integral part of a 3D printed alumina engine, possibly reinforced with carbon fibre (see also Figure 5).

44. An inner wall of engine 36. The inner wall can be an integral part of a 3D printed alumina engine (see also Figure 5). 46. A bottom of the fuel line 20 (see Figure 1) where the fuel line 20 branches off to branch pipes 28 which in turn feed the engine 36. 48. A water cooled section between the inner wall 44 and outer wall 42 which acts like a water cooled lining for the engine 36 (see also Figure 5). 50. Tops of down pipes 30 (see Figure 1) which feed H₂0₂ and/or H₂0 to the engine 36. Only one of the tops 50 is shown in Figure 4 for conciseness.

52. A wall of H₂0 tank 34 in Figure 1. The wall can be an integral part of a 3D printed alumina engine (see also Figure 5).

54. Outlets of Hydrogen and/or Methane branch pipes 28 in Figure 1. See also Figure 5.

The Bell engine exhaust 38 is not shown in Figure 2. Figure 4 shows: Only one of the outlets 54 is shown in Figure 4 for conciseness, but three outlets 54 are shown in Figure 5. The outlets 54 are spaced at 120 degrees intervals about a longitudinal axis of the craft.

56. Air flow into one of the air scoops 32. 58. Represents air scooped by scoops 32, funnelled and compressed.

60. Represents very high temperature superheated steam plus trace ΝΟχ produced by reaction between the 0₂ in the scooped air 58 and the H₂ LTACG from outlets 54. 2H₂ + 0₂ from air→ 2 H₂0 (in the form of very high temperature superheated steam + hot air (minus its Oxygen content)).

Minor ΝΟχ products_, i.e. the mono-nitrogen oxides nitric oxide and nitrogen dioxide (NO and N0₂), are produced because the extreme temperatures cause some of the atmospheric Nitrogen in the air to react with available hot Oxygen.

It is envisaged that most forms of the craft would use only atmospheric nitrogen entering the combustion chamber via the air scoops, where it would act as a chaperone molecule, increasing the efficiency of the Hydrogen Oxygen reaction. The craft 8 (indeed any flying craft in accordance with the invention) can alternatively, or additionally, comprise a nitrogen receptacle. In such a case, the nitrogen receptacle supplies nitrogen to the combustion chamber (for example 66 below).

62. H₂0₂ outlets from down pipes 30 where H₂0₂ receptacles 26 are used. Outlets 62 also shown in Figure 5. A catalyst, for example Silver oxide mesh, can be arranged across outlets 62. The catalyst initiates the break down (degradation) of the H₂0₂ into two of its component parts without chemically reacting with the catalyst.

2 H₂0₂ + catalyst→ 2 H₂0 + 0₂.

The 2H₂0 will be in the form of super heated steam from the above reaction. The 0₂ produced from the same reaction will be hot. The hot 0₂ from the breakdown of H₂0₂ can react with the H₂ provided from the Hydrogen fuel tank via outlets 54, to produce extremely superheated steam. 2 H₂ + Hot 0₂→ 2H₂0 (in the form of extremely high temperature superheated steam).

64. Ordinary steam outlets (see also Figure 5). Inner and outer walls of combustion chamber 66 described below define a space which is lined with honeycombed walls with H₂0 for cooling, due to the very high specific heat capacity of H₂0. The steam outlets 64 can be an integral part of a 3D printed alumina engine. Significant heat is produced in the H₂ + 0₂ (from air) reaction disclosed above and in the H₂ + 0₂ (from H₂0₂) reaction disclosed above. H₂0, via outlets 64, cools down the combustion chamber 66 and exhaust chamber 38 of the engine 36. The need for some steam outlets may be optional depending upon material capabilities and placement is also optional. Figures 4 and

5 show an inner group of three steam outlets 64 adjacent the three H₂0₂ outlets 62. The inner group of three steam outlets 64 are therefore adjacent the locations of the reaction of H₂ + 0₂ (from H₂0₂). Figures 4 and 5 show an outer group of three steam outlets 64 adjacent the three scooped air outlets 58. The steam outlets 64 are therefore adjacent the locations of the reaction of H₂ + 0₂ (from air). All of the steam outlets 64 may conveniently be fed by one single reservoir or a plurality of reservoirs. The same applies to steam outlets 72, 172 below. For clarity, the steam outlets 64 are venting into the combustion chamber 66 and ultimately exit into the atmosphere via the bell exhaust chamber 38.

66. Combustion chamber, for combustion of H₂ LTACG from outlets 54 with Oxygen from two sources, Oxygen from H₂0₂ source and Oxygen from air source, more specifically combustion of Oxygen from those sources which has not already reacted with the H₂ LTACG.

68. Output from Bell engine exhaust chamber 38, the output being a combination of normal steam, superheated stream, very high temperature superheated steam, extremely high temperature superheated stream, + trace ΝΟχ from the above disclosed reactions.

70. Inner and outer walls of engine exhaust chamber 38 define a space which is honeycombed with H₂0 for cooling through the wall(s) of the engine exhaust chamber. This can be an integral part of a 3D printed alumina engine.

72. Exhaust holes for steam. The exhaust holes 72 open directly to the atmosphere to allow steam to flow out and provide extra thrust (see also Figure 5). Exhaust hole 72 is one of a plurality of exhaust holes (all of which are not shown for conciseness), which could be spread evenly around the entire bell exhaust chamber wall 70 to provide a curtain of steam. The exhaust holes can be an integral part of a 3D printed alumina engine.

The wall coolant extends from the combustion chamber 66 to the exhaust chamber 38, preferably in coolant communication with the same central H₂0 reservoir which feeds outlets 64 and 72. The need for some steam outlets may be optional depending upon material capabilities and placement is also optional.

Figure 5 shows: 74. Bottom of H₂0 ballast tank 34 in Figures 1 and 4. This can be an integral part of a 3D printed alumina engine.

76. Inner wall of bell exhaust chamber. In the embodiment shown, the wall is circular. This can be an integral part of a 3D printed alumina engine.

78. Outer wall of bell exhaust chamber. In the embodiment shown, the wall is circular. This can be an integral part of a 3D printed alumina engine, possibly reinforced

with carbon fibre.

In more detail, the engine of the Potent Voyager craft 8 works as follows :-

Because the craft 8 is lighter than air, the passive lift initially causes movement upwards, forcing air into the air scoops 32, without the need for moving parts. The air scoops 32 of the craft 8 funnel and compress the air at various angles into the engine 36. When Hydrogen 18 is released into the combustion chamber 38, there are a number of options to ignite the Hydrogen in the combustion chamber.

1) H₂0₂, which has been ignited by a catalyst, is then used to ignite the Hydrogen.

2) If the pressure of the Hydrogen is greater or equal to 10 bar, or if the pressure differential between the Hydrogen 18 and the gas(es) outside the craft is high enough, the Hydrogen will self ignite in the presence of Oxygen in the air in the combustion chamber. 3) Other existing ignition systems could be chosen and added to our basic system depending upon the flight profile required.

4) For continual ignition once the craft has reached sufficient velocity, the compression and friction of the air as it enters the combustion chamber via the air scoops, heats the air with sufficient levels of thermal properties to combust with the Hydrogen.

The engine 36 will basically be a cylinder with holes in the sides to let in the air from the air scoops

32 and pipes 28 to let in the Hydrogen. The outer conical surface of the receptacle 34 is at an angle of 60-70 degrees to the horizontal. Moreover, the angle of the outer conical surface of the receptacle 34 to the horizontal is conveniently about 1-2 degrees different from the angle of the air scoop to the horizontal so as to encourage turbulent fuel mixing. The ballast water in the receptacle 34 is used to cool the engine. The ballast water in the receptacle 34 is also a source of additional thrust, when the engine turns ballast water into high pressure steam it is discharged via downward facing pipes 64 and 72. H₂0₂ could also be used in the engine 36 in a similar manner. That is more akin to a traditional rocket engine. This is part of the hybrid modes of the engine 36, which can be used alongside the air breathing mode, or as purely in a rocket mode when in hard vacuum, if additional valves are used to close off the air intakes and the scooped air flow 58, once the gaseous fuels H₂/CH₄ have been expended. Also in the rocket mode if the gaseous fuels have not all been expended, this can also be used in the engine to provide thrust.

The flexibility of the engine 36 has many operational permutations of which the primary modes of operation are listed below. 1) Air breathing lighter than air combustible gases.

2) Air breathing lighter than air combustible gases and super heated steam from the water coolant ballast. 3) Air breathing lighter than air combustible gases and super heated steam from the water coolant ballast and superheated steam from the H₂0₂ coolant ballast.

4) Rocket thrust from H₂0₂ only ( for use in vacuum ). 5) Lighter than air combustible gases using the Oxygen from the H₂0₂ reaction in addition to the super heated steam from that reaction. ( for use in vacuum ).

6) Lighter than air combustible gases using the Oxygen from the H₂0₂ reaction in addition to the super heated steam from that reaction and the super heated steam from the water coolant ballast ( for use in vacuum ).

7) The lighter than air combustible gases as a purely gaseous thrust (for use in vacuum). 8) Steam from any remaining water coolant ballast because it would naturally boil off in a vacuum ( for use in vacuum ).

Summary of inputs, reactions, and outputs, for the Potent Voyager class craft 8 of Figures 1 to 5, with particular reference to Figure 4 :

Inputs for fuels reactions both chemical and physical are Gaseous Hydrogen (¾), Atmospheric Air (Mostly Nitrogen (N₂) and Oxygen (0₂)), Rocket grade Hydrogen Peroxide (H₂0₂), and liquid water (H₂0).

Outputs during atmospheric flight, the H₂ will react with the 0₂ in the air to produce very superheated steam as one of the exiting gases. There may also be minor amounts of ΝΟχ gases produced depending upon the mixtures used. Outputs during atmospheric flight and vacuum flight from H₂ and H₂0₂ will be superheated steam from the H₂0₂ reaction disassociating into superheated steam and superheated 0₂ which could then react with some of the H₂ to produce extremely high temperature superheated steam, again minor amounts of ΝΟχ could be produced depending upon the mixtures used. Outputs during atmospheric flight and vacuum flight from liquid H₂0 will be turned into steam via cooling of both air scoops and engine reducing the need to use additional protective mass on both items. This involves a phase change of states of the H₂0 from liquid to gaseous form and not a chemical reaction. Waste heat from the other reactions is used to do this. Use of atmospheric N₂ as reaction mass as it is heated by the air scoops and the reactions in the engines. This is where some of the minor ΝΟχ emissions may arise, but most of the N₂ will be expelled unreacted, just at a hotter temperature.

As the LTACG is combusted, it acts as a ballast, increasing the buoyancy at the level of dumping.

ISP's from the various fuels used.

Traditionally liquid Oxygen (LOX) and liquid Hydrogen (LH₂) give an ISP of 455 seconds in a vacuum whilst within the earth's gravitational field before orbital velocity has been obtained, but by using non cryogenic gaseous forms of both gases, it would considerably increase ISP purely on a chemical reaction efficiency basis, not having to warm both fuels from cryogenic temperatures. In addition to this it should also gain ISP efficiency by not carrying the atmospheric 0₂, and also by not having to waste energy lifting any of the fuels whilst they are displacing their weight in air.

ISP of H₂0₂ and gaseous H₂ should be at least 350 seconds not counting much greater efficiency whilst displacing atmospheric air.

ISP of turning liquid H₂0 in to gaseous steam H₂0 is traditionally 190 seconds, again not counting much greater efficiency whilst displacing atmospheric air.

The optimal places for the water cooling of both the engine and air scoops will depend upon the amount of mass used in construction of both, and using some of it in the engine could increase overall ISP. Hence the possibility of using it in the side walls of the engine or both with separate consideration taken for the air-scoops depending upon the size used.

The combination of all these different ISP's at different times should combine to give the ability to produce a SSTO (Single Stage To Orbit) hybrid vehicle with an average ISP of greater than 600- 700 seconds, a theoretical milestone not previously thought of that it could be done chemically, other than possibly with an air breathing vehicle.

Further calculations have shown whilst lighter than air:-

Waste heat dumped into H₂0 gives an approximate ISP of 1900 seconds.

H₂0₂ by itself gives an approximate ISP of 1610 seconds + when used with H₂ gives an approximate ISP of 3519 seconds at 1 atmospheric pressure.

H₂ if burnt with carried 0₂ gives an approximate ISP of 3997 seconds at 1 atmospheric pressure.

H₂ if burnt with atmospheric air gives an approximate theoretical maximum ISP of 80,000 seconds at around Mach 5 to 6. When water is heated to well over 2000 °C, a small percentage of it will decompose into OH, monatomic Oxygen, monatomic Hydrogen, 02, and H₂. If the engines are run at these temperatures, there may be some additional ISP thrust. This is an endothermic process which should lower the amount of cooling needed to prevent the engine walls from melting.

Referring to Figure 6, a so-called Coanda Class of launch craft 98 is to be used as human safe launch vehicles.

The Coanda class of craft 98 will launch in an upright stance with a gaseous passive/active lift fuel at the top, with a passive neutral lift gas as part of the safety separation outer layer, followed by a concentric ring level going from outer to inner in the order of :- - Human safe payload area with life support,

- With the interchangeable and differing variables of water ballast, coolant and or fuel,

- Air/strut gap and then engine in the central point of the Coanda vehicle,

- With the interchangeable & differing variables of H₂0₂ ballast if any, coolant and/or fuel, and

- Other payload.

The craft 98 looks like an elongated tubular ring doughnut designed to take advantage of the Coanda effect (i.e. that a moving stream of fluid in contact with a curved surface will tend to follow the curvature of the surface rather than continue travelling in a straight line) to suck and accelerate a mass of air around and in through the centre of the launch craft 98, then into and around the engine 138, in addition to the thrust from the rocketry effect. Atmospheric air is therefore part of the reactive mass of the craft, i.e. the fuel shot backwards from the craft to provide propulsion.

The passive lift of the vehicle initiates the Coanda effect, causing the air to flow around the vehicle and be directed into and around the engine. The Coanda engine uses the Coanda effect to draw atmospheric air into the engine at an accelerated rate, to be used in the combustion cycle of the engine, by the reactive combustion of lighter than air gas or gases, (Hydrogen, Methane), and the ballast coolant or coolants (H₂0 or H₂0₂, or both), to produce thrust in the form of superheated steam. This can be done passively and or actively or both. Other aspects of the engine are very similar to the Potent Voyager engine but with the air and gaseous fuel intakes swapped around. With the air scoops 151 centrally located as in a topological inversion of the Potent Voyager engine

36. There are at least the same number of operational modes as that of the Potent Voyager engine 36 listed above. Figure 6 shows:

The craft 98.

100. Indicating the ring shape of the air intake for the Coanda launch craft 98.

102. Re-enforcement analogous to reinforcement 10 in Figure 1.

104. An outer envelope analogous to reinforcement 12 in Figure 1. 106. An inner envelope analogous to inner envelope 14 in Figure 1.

108. Helium gas analogous to Helium gas 16 in Figure 1.

110. Receptacle for a primary lighter than air gaseous fuel analogous to primary lighter than air gaseous fuel 18 in Figure 1.

111. A payloads module. This module structure contains the payload and ballast/secondary fuel tanks (which can be made from H.D.P.E.), and can be constructed from carbon fibre, which can be glued and/or welded to the outer envelope 104. 112. A top of the payloads module 111. The inner side of the top 112 of the payloads module 11 1 connects directly to the lamina flow struts 115, which connect directly to the Coanda engine 138. There are spaces between the lamina flow struts 115 for air to pass from the inner side of the top 112 of payload module 11 1, down, and by pass the engine 138. Air scoops 151 , described below, are situated on top of, and as part of, the engine 138 also described below. These features are not necessarily to scale with the rest of the module 111 in the Figure 6 representation. See also Figures 9 and 10. 114. One of three fuel lines for primary lighter than air gaseous fuel H₂ &/or CH₄ displaced by pressure differentials into the Coanda engine. This can be constructed from H.D.P.E. (High-density polyethylene), (see also Figure 7, but the full extent of the fuel lines 114 is shown only in Figure 6).

115. A lamina flow strut (there are three in the embodiment shown). See also Figures 7 to 10.

116. Edge on lamina flow strut connected to the top of the payload module. This can be made of carbon fibre.

118. Leading edge of lamina flow strut. This can be made of carbon fibre, (see also Figure 7, 8 and 9)·

136. Trailing edge of lamina flow strut. This can be made of carbon fibre. See also Figure 7 and Figures 9 to 10.

The lamina flow struts contain the pipework 114 of H₂ &/or CH₄, H₂0₂, and H₂0 that are used to feed the engine.

The laminar flow strut outer surface is symmetrical about a line through the leading edge 118 and the trailing edge 136, so as to ensure no net lifting forces applied transverse to the surface of the struts.

The struts can also be configured at different angles to control airflow and craft orientation. The number of struts can also be increased for more control and structural strength depending on the flight profile required.

120. Pressurised helium above H20 ballast tank used to pump H20 into engine. In a fixed volume container a pressurised gas e.g. helium can push a liquid or gas also in said container out of said container through a valve system without the need for active pumps.

122. Pressurised helium above H₂0₂ ballast tank used to pump H₂0₂ into engine.

124. Human safe escape capsule with fire proof parachute. This can be made of carbon fibre, (see also Figure 7). 126. Payload (crew and/or passengers and/or cargo) with fire proof parachute. Figure 7 shows three payloads spaced at 120 degrees about a longitudinal axis of the craft. 128. H₂0₂ ballast tank. This can be constructed from H.D.P.E. (High-density polyethylene). There is pressurised helium gas above the H₂0₂ ballast in the tank 128 used to pump H₂0₂ into engine 138 although pressurised helium gas is not shown for conciseness, (see also Figure 7).

130. H₂0 ballast tank. This can be constructed from H.D.P.E. (High-density polyethylene). There is pressurised helium gas above the H₂0 ballast in the tank 130 used to pump H₂0 into engine 138 although pressurised helium gas is not shown for conciseness, (see also Figure 7).

Features 128 and 130 are interchangeable features that can also be used in combination. Features 120, 122, 128, 130 may be optional depending on the flight profile required.

132. H₂0₂ Pressurised pipe to feed coanda engine. This can be constructed from H.D.P.E. (High- density polyethylene) connected to alumina pipework part of a 3D printed alumina engine, (see also Figure 7).

134. H₂0 pressurised pipe to feed coanda engine. This can be constructed from H.D.P.E. (High- density polyethylene) connected to alumina pipework part of a 3D printed alumina engine, (see also Figure 7). The H₂0₂ pipework 132 and the H₂0 pipework 134 represented by the dotted lines refer to the pipework from ballast tanks obscured by the ballast tanks numbered 128 and 130.

138. Coanda engine, (see also Figure 7). Expanded details on Figures 8, 9 and 10. This can be a 3D printed alumina engine.

140. Indicating the ring shape of the trailing end of the Coanda launch craft for optimal air flow utilising the Coanda effect. It should be appreciated between the ring shape air intake 100 and the ring shape of the trailing end 140, the outer envelope 104 has a continuous ring shape. It will also be noted the distance between inner facing surfaces of the outer envelope 104 decrease towards the mid length of the outer envelope 104, and increase again towards the trailing end of the outer envelope 104.

Figure 7 shows:

142. Living area of Coanda control module. The living area comprises air at Standard temperature and pressure (STP). The STP air 142 surrounds all of the modules 124, 146, 126, 128, and 130.

144. (Circular) inner side of Coanda module corresponding to number 112 on Figure 6. This can be made of carbon fibre. The inner side of Coanda module 144 connects directly onto the lamina flow struts 1 15, which connect directly onto the Coanda engine 138. This area between the inner side of the Coanda module 144 is not necessarily to scale with the outer radius of the module in this representation. The position of the inner side 144 of Coanda module is schematically represented in Figure 8. The point on the inner side of the Coanda module referenced 144 in Figure 8 is slightly higher than the point on the inner side of the Coanda module referenced 144 in Figure 7, See also Figure 10.

146. Life support machinery to maintain a habitable environment. Life support machinery 146 is in gaseous communication (i.e. breathable air) with 124 and/or 126 and/or throughout living area 142.

148. Outer wall of life support, payload and operational control module. This can be made of carbon fibre.

Figure 8 shows an expanded scale plan view of the engine 138 in relation to the inner side 144 of the Coanda module. The scale of the gap between the inner side 144 of the Coanda module and the engine 138 has been compressed. In practice, the inner side 144 of the Coanda module would be much further out from the engine 138.

Other parts shown in Figure 8 :

118. Leading edge of lamina flow strut 115. This can be made of carbon fibre. See also Figure 9. The leading edge 118 of lamina flow strut inclines at an angle of between 60-70 degrees to a horizontal line across the sheet. The trailing edge 136 of lamina flow strut inclines at an angle of about 30 degrees to a horizontal line across the sheet. See also Figure 9.

151. Three air scoops. In other embodiments (not shown), a different number of a plurality of scoops is envisaged. It will be appreciated that the craft is designed so that during flight a quantity of air passed through the gap between the inner side of Coanda module 144 and the outer edge of the scoops 151.

150. Outer wall top of the three air scoops 151 shown as an integral part of the top of the engine 138. This feature can be an integral part of a 3D printed alumina engine, possibly reinforced with carbon fibre.

156. Inner side walls of one the air scoops. This feature can be an integral part of a 3D printed alumina engine, possibly reinforced with carbon fibre. The inner side walls each incline downwardly, optimally at an angle of between 60-70 degrees to the plane of the sheet, towards the air inlet described below. The three side walls 156 cooperate to form a conical (in other words a funnel shaped) air scoop 151 with a narrowing at the bottom of the conical scoop adjacent the air inlet 152. The inner side walls 156 are H₂0 cooled as described below. 152. Air inlet into the engine 138 for compressed atmospheric air compressed by outer wall of air scoop 150.

Features 152 and 156 can be an integral part of a 3D printed alumina engine. The air scoops 151 are located above the engine 138. The air scoops 151 compress the air flow directly into the combustion chamber (180 in Figure 9) of the engine 138 from above. The Bell engine exhaust chamber (181 in Figure 9) is not shown in Figure 8 but it would be located between the outer wall top 150 of the three air scoops 151 and the inner side of the Coanda module numbered 144.

154. Top of H₂0 ballast tank cone. See also Figure 9. The scale of the engine 138 in Figure 6 is too small to show much detail, but the triangular shaped object on the top of the engine 138 represents a possible position of the top 154 of H₂0 ballast tank cone. This feature can be an integral part of a

3D printed alumina engine, possibly reinforced with carbon fibre. Figure 9 shows a 180 degree cross-section. Due to the spacing of the air-scoops, it is useful to show just one air-scoop, air inlet 152, and corresponding pipe work (on the left hand side of Figure 9) and just one laminar flow strut 115 (on the right hand side of Figure 9).

114. Fuel line.

118. Leading edge of lamina flow strut 115. 132. Η₂0₂ Pressurised pipe to feed Coanda engine 138. 134. H₂0 pressurised pipe to feed Coanda engine. 150. Outer wall top of one of the air scoops 151. 152. Air inlets.

154. Top of H₂0 ballast tank cone.

156. H₂0 cooled inner side walls of air scoops.

158. Helium gas pipe for cooling of the leading edge of the lamina flow strut 115. This pipe can be constructed from High-density polyethylene ( H.D.P.E.) connected to alumina pipework that is part of a 3D printed alumina engine.

160. Air flowing into air scoop 151.

162. Air compressed in air scoop 151.

164. Steam produced by the cooling of the engine 138 and recycling heat produced by drag. Flow path to steam outlets 172 disclosed below.

166. Helium gas pipe for cooling of the trailing edge of the lamina flow strut 115. This can be an integral part of a 3D printed alumina engine. 136. Trailing edge of lamina flow strut.

170. Outlet for primary lighter than air gaseous fuel (H₂ and/or CH₄) displaced by pressure differentials into the engine. The outlet can be an integral part of a 3D printed alumina engine. An appropriate safety valve will be placed inline upstream of the outlet 170 (See also Figure 10) See Figures. 11 and 12 for possible examples of safety valve.

172. Ordinary steam outlets. Figure 9 shows an inner group of outlets and an outer group of outlets. Figure 10 shows three inner outlets 172. Figure 9 shows an outer outlet associated with each air scoop, implying three outer outlets 172. The outlets can be an integral part of a 3D printed alumina engine. Placement and need for some ordinary steam outlets may be optional depending upon material capabilities. Referring to Figures 13 to 16, each air scoop 151 has an outer wall coolant flow system 171 (outer wall) and an inner wall coolant flow system 171 (inner wall) associated with it.

The outer wall coolant flow system 171 (outer wall) has a first part (see Figure 13 and 16) and a second part (see Figure 14 and 16). Referring to Figure 13 and 16, the outer wall coolant flow system 171 (outer wall) first part, for one air scoop 151, for simplicity in this example, has inlets

171, from one of the steam chambers/water ballast tanks 164. The first outer wall coolant flow system 171 (outer wall) spirals down (in practice, more tightly and with more loops), as pipe work embedded in the walls of the engine, to the part of the scoop 151 adjacent the openings 152. Referring to Figures 14 and 16, the outer wall coolant flow system 171 (outer wall) second part zig zags as it travels from adjacent the opening 152 to one of the outer outlets 172 associated with each air scoop 151.

Referring to Figures 15 and 16, an inner wall coolant flow system has three channels, one associated with each air scoop 151. Each channel has an inlet from one of the steam chambers/water ballast tanks 164 underlying the H₂0 cooled inner side wall 156 of its associated air scoop 151 nearest the top 154 of H₂0 ballast tank cone. Referring to Figure 15 in particular, each channel spirals down, as pipe work embedded in the walls of the engine, to one of the inner outlets 172 (shown also in Figure 16) associated with each air scoop 151. Figure 16 shows the outer wall coolant flow system and the inner wall coolant flow system.

Referring back to Figure 9: 174. Air 160 + H₂ (an example of a primary lighter than air gaseous fuel) exiting outlet 170 produces very high temperature superheated steam + trace ΝΟχ, i.e. the mono-nitrogen oxides nitric oxide and nitrogen dioxide (NO and N0₂).

176. A holed disc (or a ring - not shown for conciseness) in communication with angled H₂0₂ pressurised pipe 132. The holed disc (or ring) functions as a delivery system for H₂0₂ into the combustion chamber of the engine, and the subsequent reaction, the reaction being initiated by a silver oxide catalyst gauze mesh over the holes in the holed pipe 176. The holed pipe 176 can be an integral part of a 3D printed alumina engine. The silver oxide catalyst at the H₂0₂ outlet causes the H₂0₂ from the holed disc (or ring) 176 and the H₂ from outlet 170 to turn into superheated steam (H₂0) + superheated 02 referenced 178. The H₂ from outlet 170 can then react with the superheated 02 178 to produce extremely high temperature superheated steam as disclosed hereunder, in an a reaction similar to the reaction described in relation to H₂0₂ outlets 62 above.

180. Combustion chamber.

181. Bell exhaust chamber. 182. Output of Bell exhaust chamber 181. The output 182 is the average of normal steam (from steam outlets 172), superheated steam (178 from the reaction of the H₂0₂ and the H₂ catalytic reaction), very high temperature superheated steam (174 from the reaction of air 160 and H₂), and extremely high temperature superheated steam (from the reaction of superheated 02 178 and H₂ from outlet 170, and trace ΝΟχ (from the nitrogen in the compressed atmospheric air, which can react with either the Oxygen in the compressed air or the Oxygen produce by the catalytic reaction of the H₂0₂) from the above disclosed reactions.

If CH₄ LTACG is used instead of H₂, superheated C0₂ will also be produced in addition to the products of the reactions above. 184. One of a plurality of exhaust holes for steam cooling through the walls of exhaust chamber 181 of the engine 138. The steam flowing through the exhaust holes can provide extra thrust. The exhaust holes are in a configuration spread evenly around the entire bell exhaust chamber wall to provide a curtain of steam. This can be an integral part of a 3D printed alumina engine. Placement and need for some steam outlets maybe optional depending upon material capabilities.

Figure 10 shows: 152. Air inlet.

136. Trailing edge of lamina flow strut 115.

170. Outlet for primary lighter than air gaseous fuel (H₂ and/or CH₄) displaced by pressure differentials into the engine 138.

Appropriate safety valve examples of which are discussed below in connection with Figures 11 and 12 can be placed inline upstream of the outlet 170. 172. Ordinary steam outlets. Placement and need for some steam outlets maybe optional depending upon material capabilities. Not all ordinary steam outlets are shown.

176. A holed disc (or ring). 184. Exhaust holes.

Referring to Figure 16, in this example, for simplicity, each of the steam chambers/water ballast tanks 164's is connected to inlets 183, from which point flow channels tightly spiral down, as embedded pipe work in the walls of the engine, to three outlets 184.

Other arrangements for the (outlets 172 and) exhaust holes 184 of the water/steam cooling system for the engine, and extra thrust, are envisaged.

186. Wall of bell exhaust chamber. This can be an integral part of a 3D printed alumina engine. 188. Wall of combustion chamber. This can be an integral part of a 3D printed alumina engine.

190. Bottom of H₂0 ballast tank cone. This can be an integral part of a 3D printed alumina engine. The scale of the engine 138 in Figure 6 is too small to show much detail, but the triangular shaped object below the engine 138 represents a possible position of the bottom 190 of H₂0 ballast tank cone.

192. Three inlets for H₂0 ballast tank cone (only one of which is referenced for conciseness). This feature would not be seen from below but is shown in Figure 10 to show position of the inlets corresponding to pipework 134 in Figure 9 The inlets 192 can be an integral part of a 3D printed alumina engine.

194. Three inlets (only one inlet is referenced for conciseness) for H₂0₂ holed disc (or ring). This feature would not be seen from below but is shown in Figure 10 to show position of inlets 194 corresponding to pipework 132 in Figure 9. The inlets 194 can be an integral part of a 3D printed alumina engine.

Summary of inputs, reactions, and outputs, for the Coanda class craft 98 of Figures 6 to 10, with particular reference to Figure 9:

Inputs for fuels reactions both chemical and physical are Gaseous Hydrogen (H₂), Atmospheric Air (Mostly Nitrogen (N2) and Oxygen (02)), Rocket grade Hydrogen Peroxide (H₂0₂), and liquid water (H₂0).

Outputs during atmospheric flight, the H₂ will react with the 02 in the air to produce very superheated steam as one of the exiting gases, there may also be minor amounts of NO_x gases produced depending upon the mixtures used. Outputs during atmospheric flight and vacuum flight from H₂ and H₂0₂ will be superheated steam from the H₂0₂ reaction disassociating into superheated steam and superheated 02 which could then react with some of the H₂ to produce extremely high temperature superheated steam again minor amounts of ΝΟχ could be produced depending upon the mixtures used. Outputs during atmospheric fiight and vacuum flight from liquid H₂0 will be turned into steam via cooling of both air scoops and engine reducing the need to use additional protective mass on both items. This involves a phase change of states of the H₂0 from liquid to gaseous form and not a chemical reaction, waste heat from the other reactions is used to do this.

Use of atmospheric N2 as reaction mass as it is heated by the air scoops and the reactions in the engines. This is where some of the minor ΝΟχ emissions may arise, but most of the N2 will be expelled unreacted, just at a hotter temperature. ISP's from the various fuels used.

Traditionally LOX and LH₂ give an ISP of 455 seconds in a vacuum whilst within the earth's gravitational field before orbital velocity has been obtained, but by using non cryogenic gaseous forms of both gases, it would considerably increase ISP purely on a chemical reaction efficiency basis, not having to warm both fuels from cryogenic temperatures. In addition to this it should also gain ISP efficiency by not carrying the atmospheric 02, and also by not having to waste energy lifting any of the fuels whilst they are displacing their weight in air.

ISP of turning liquid H₂0 in to gaseous steam H20 is traditionally 190 seconds, again not counting much greater efficiency whilst displacing atmospheric air. The optimal places for the water cooling of both the engine and air scoops will depend upon the amount of mass used in construction of both, and using some of it in the engine could increase overall ISP. Hence the possibility of using it in the side walls of the engine or both with separate consideration taken for the air-scoops depending upon the size used. The combination of all these different ISP's at different times should combine to give the ability to produce a SSTO (Single Stage To Orbit) hybrid vehicle with an average ISP of greater than 600- 700 seconds, a theoretical milestone not previously thought of that it could be done chemically, other than possibly with an air breathing vehicle. Further calculations have shown whilst lighter than air:-

Waste heat dumped into H₂0 gives an approximate ISP of 1900 seconds.

H₂ if burnt with carried 0₂gives an approximate ISP of 3997 seconds at 1 atmospheric pressure.

H₂ if burnt with atmospheric air gives an approximate theoretical maximum ISP of 80,000 seconds at around Mach 5 to 6.

When water is heated to well over 2000 °C, a small percentage of it will decompose into OH, monatomic Oxygen, monatomic Hydrogen, 0₂, and H₂. If the engines are run at these temperatures, there may be some additional ISP thrust.

This is an endothermic process which should lower the amount of cooling needed to prevent the engine walls from melting.

Some of the atmospheric air will be used as a reaction mass by the flying wing aspect of the vehicle as a whole, due to the Coanda/Bernoulli/Newton effect.

Figure 11 shows:

A version 1 of a safety valve 24, 198. 200. Primary LTACG (H₂ and/or CH₄).

202. Gas impermeable wall of safety valve pipe. This can be an integral part of a 3D printed alumina engine.

204. One of a plurality of concave sections such as a parabolas (only one being referenced for conciseness) in gas impermeable wall of safety valve pipe. The concave sections 204 can be an integral part of a 3D printed alumina engine. The concave sections 204 can be used to deflect and interfere with shock waves of a back blast from the combustion chamber 66 of the engine 36, 138.

206. Focal points of concave sections/parabolas (only one being referenced for conciseness).

208. Expansion points in the pipework. These features can be an integral part of a 3D printed alumina engine.

210. H₂0 reservoir to isolate LTACG 200 from back blasted oxidising agents. The H₂ LTACG 200 passes safely through the H₂0 210, in the form of bubbles, because the H₂0 is also saturated with H₂.

212. Pipe restriction to increase fuel gas pressure into engine. The pipe restriction can be an integral part of a 3D printed alumina engine. The pipe restriction also reduces back blast gas pressure from the combustion chamber 66 of the engine 36, 138.

214. LTACG 200 exit towards combustion chamber 66 of engine 36, 138.

Figure 12 shows a version 2 of a safety valve 24, 218 for supersonic gaseous fuel injection:

220. Primary LTACG (H₂ and/or CH₄).

222. Gas impermeable wall of safety valve pipe. This can be an integral part of a 3D printed alumina engine.

224. Parabola in gas impermeable wall of safety valve pipe used to reflect back blast from combustion chamber from the engine. This can be an integral part of a 3D printed alumina engine.

226. Focal points of parabola.

228. Pipe restriction to increase fuel gas pressure into engine and reduce back blast gas pressure. 230. Top pressure cap. This can be made from disposable metal foil. 232. Pipe work constricted to hourglass shape. This can be an integral part of a 3D printed alumina engine.

234. Vacuum pipe, exact length will need to be calculated for specific profiles.

236. Bottom pressure cap. This can be made from disposable metal foil.

238. LTACG exit towards combustion chamber of engine 36, 138. Both engines 36, 138 are capable of hybrid ballast use which can be used as a coolant and extra thrust. This can be H₂0 and/or H₂0₂, but fuels other than H₂0₂ could also be used.

Both engines 36, 138 work as follows: For a 100% burn a minimum ratio of 2.5 volumes of air to 1 volume of gaseous Hydrogen needs to be achieved in the reaction chamber of the engine. The air scoops need to be scaled appropriately, to achieve this constantly, along a column of air for the chosen ascent profile. In reality, a lean burn ascent profile (greater volume of air) gives more flexibility in reaching these proportions, with the excess mass of air being expelled as a reaction mass, and therefore giving the possible ability to reach the higher theoretical maximum ISP of gaseous Hydrogen.

Hybrid Ram-Pulse- Jet combustion cycle

Air is compressed into the engine initially via passive lift. Gaseous Hydrogen fuel is injected into the engine by both pressure differential and Coanda effect of the compressed air going into the engine. The nearly parallel flows of these two gaseous forms are turbulently mixed, this is then ignited. The expanded superheated steam mostly exits the nozzle, providing active thrust via the exhaust. The back-blast suppresses the next combustion cycle until sufficient fuel and air (the air now having active compression in addition to the passive lift compression) enters the combustion chamber for the next cycle. This is initially pulse jet like, but at sufficient velocity it should act more like a continuous ramjet.

Rather than having weighty and complex turbo pumps to inject the gaseous fuel into the engine, it will be using pressure differentials to inject the gaseous fuels into the engine. As the balloons are fixed volume, the higher they go, the more pressure difference there is between the inside and the outside. As pressure drops, the Hydrogen will be forced out through the bottom of the balloon connected to the engine pipes going through a water buffer (sufficient to prevent the engine from melting, but ideally the last few drops will be turned into steam as the last wisps of the Hydrogen are burnt. Later, larger versions may carry excess water as water can be useful on manned missions) acting as a safety buffer and to pre -heat the Hydrogen, as warm Hydrogen will burn better (more ISP) than cold. This will be another advantage it has over cryogenic Hydrogen and LOX.

A further advantage could be obtained by over -pressuring the gaseous fuel, so that the fuel is supersonically injected into the combustion chamber. This is done by correctly configured hourglass shapes in the pipework. This could be improved with a vacuum chamber placed further along in the pipework, and used in a similar manner to the chambers used in a traditional pulse jet cycle, to create a supersonic injection cycle. Thereby giving a greater thrust by adding the supersonic velocity of the un-combusted fuel to the supersonic speeds obtained by the combustion reaction, as it is expelled through the exhaust nozzle (see Figure 12).

Another advantage of the system is that, like the mouse and elephant heat loss analogy, it's more efficient the larger it is, as will be the surface to volume ratio. As the surface to volume ratio is more beneficial the larger it is, in this case. This is because less surface area mass of balloon envelope is used proportionally to the volume of primary lighter than air gaseous fuel. Therefore giving a better usable energy to mass ratio, and an increase in the choice of the amount of passive lift available.

Yet another advantage the applicant's craft has over traditional rocketry is that it does not expend oxidiser and fuel whilst in pre launch mode. This is because it does not have to vent the boil off of the cryogenic fuels to prevent catastrophic fails from pressure build up.

Referring to Figures 17 to 21, the Liberator crafts 298, 308, 318 are a heavy lift class of craft comprising (three or as shown)four modules of the Potent Voyager class craft 8 and/or the Coanda class craft 98. Figure 17 and 20 show:

A craft 298.

300. Potent Voyager Launch craft units. Payload sections can be removed. 302. Lamina flow wings corresponding to the Potent Voyager class craft. This can be made from carbon fibre. 304. Potent Voyager Launch craft unit. Payload section can be expanded.

The dot at the centre of each circle 300, 304 denote the top/pinnacle of the reinforced section 10. The outer circumference of each circle 300, 304 corresponds to the outer envelope 12 of each Potent Voyager craft. Mass/payload (including that of the craft itself) is distributed throughout the craft 298 to ensure the craft is balanced, and does not topple. The mass of the payload may generally be located in the bottom unit 304 of the craft 298 in the configuration shown in Figures 17 and 20. The centre of mass of the craft 298 acts substantially in the vertical line through the centres of the two lowest units 302, 204 of the craft 298.

The three upper most lamina flow wings 302 in Figures 17 and 20 can be asymmetrical/aerofoil shaped. Such an arrangement can provide lift to the craft 298 in the configuration shown in Figures 17 and 20. The lowest lamina flow wing 302 in Figures 17 and 20 can be symmetrical. Such an arrangement ensures the craft 298 is symmetrical, in terms of wing induced lift forces, about a vertical line through the centres of the two lowest units 302, 204 of the craft 298. The arrangement can benefit the anti roll properties of the craft 298. However, other arrangements are envisaged that do not require wings 302 of the type mentioned hereinbefore.

In use, the craft 298 take off horizontally e.g. from a mast or from an open patch of ground, No runway or aeroplane type landing gear is required.. One or more of the Potent Voyager components 300, 304 can be rotated through 90 degrees to provide thrust for horizontal take off of the craft 298. In such a case, some pipework and H₂0 & H₂0₂ tanks will need to be rotated by 90 degrees in relation to their position on Figures 1-5. Immediately after take off the craft 298 acts like a cross between an airship and aeroplane. The pitch of the craft 298 increases as it ascends, to get out of the atmosphere, but the pitch does not reach the vertical. Once the craft has substantially passed the atmosphere, the pitch of the craft 298 decreases, so the craft can make a horizontal/tangential orbit. It is not necessary to change the shape of the asymmetrical lamina flow wings 302, during the flight, as, once out of the atmosphere, they do not provide significant lift, and the rocket aspect would provide the main directional thrust. Traditional attitude jets can provide fine orientation control of the craft 298 as needed. It will be appreciated that the flight trajectory of the craft 298 takes advantage of the atmosphere, using the atmosphere for air-breathing and wing induced lift.

Figures 18 and 21 show: A craft 308.

310. Coanda Launch craft units with payload sections removed. 312. Lamina flow wings Coanda class. These can be made from carbon fibre. 314. Coanda Launch craft unit with expanded payload section.

The circular parts of the Coanda craft 310, 314 referenced in Figure 18 corresponds to the outer envelope 104 and the inside edge of the payload module 112. Within the circular parts of the Coanda craft 310, 314 referenced in Figure 18 the three lines spaced at 120 degrees correspond to the laminar flow struts 115.

Mass/payload (including that of the craft itself) is distributed throughout the craft 308 to ensure the craft is balanced, and does not topple. The mass of the payload may generally be located in the bottom unit 314 of the craft 308 in the configuration shown in Figures 18 and 21. The centre of mass of the craft 308 acts substantially in the vertical line through the centres of the two lowest units 310, 314 of the craft 308. The three upper most lamina flow wings 312 in Figures 18 and 21 can be asymmetrical/aerofoil shaped. Such an arrangement can provide lift to the craft 308 in the configuration shown in Figures 18 and 21. The lowest lamina flow wing 312 in Figures 18 and 21 can be symmetrical. Such an arrangement ensures the craft 308 is symmetrical, in terms of wing induced lift forces, about a vertical line through the centres of the two lowest units 312, 314 of the craft 308. The arrangement can benefit the anti roll properties of the craft 308. However, other arrangements are envisaged that do not require wings 312 of the type mentioned hereinbefore.

In use, the craft 308 take off horizontally e.g. from a mast or from an open patch of ground, No runway or aeroplane type landing gear is required.. One or more of the Potent Voyager components 310, 314 can be rotated through 90 degrees to provide thrust for horizontal take off of the craft 308. In such a case, some pipework and H₂0 & H₂0₂ tanks will need to be rotated by 90 degrees in relation to their position on Figures 6-10. Immediately after take off the craft 308 acts like a cross between an airship and aeroplane. The pitch of the craft 308 increases as it ascends, to get out of the atmosphere, but the pitch does not reach the vertical. Once the craft has substantially passed the atmosphere, the pitch of the craft 308 decreases, so the craft can make a horizontal/tangential orbit. It is not necessary to change the shape of the asymmetrical lamina flow wings 312, during the flight, as, once out of the atmosphere, they do not provide significant lift, and the rocket aspect would provide the main directional thrust. Traditional attitude jets can provide fine orientation control of the craft 308 as needed. It will be appreciated that the flight trajectory of the craft 308 takes advantage of the atmosphere, using the atmosphere for air-breathing and wing induced lift.

Figure 19 shows: A craft 318.

300. Potent Voyager Launch craft units with payload sections removed.

310. Coanda Launch craft unit with payload sections removed.

320. Lamina flow wings combined class. This can be made from carbon fibre.

304. Potent Voyager Launch craft unit with expanded payload section. The craft 318 is arranged in a similar manner to the crafts 298, 308, and follows a similar flight trajectory.

The Liberator type craft 298, 308, 318 uses Hydrogen peroxide as the coolant for the engine for lifting heavy loads, both to L.E.O. and higher, to give extra thrust and the ability to burn excess Hydrogen when out of the atmosphere and in hard vacuum. The launch profile will require a longer, slower ascent profile whilst in atmosphere to achieve eventual escape velocity. For much larger payloads it will need to continue to accelerate when out of the atmosphere by the combustion of the HTP ballast and the primary fuel gases. The Liberator type craft 298, 308, 318 may be better for certain angled satellite launches, with rapid reuse of the first stage possible, for heavy bulk lifts into space.

The buoyancy of the crafts 8, 98, 298, 308, 318 is controllable. It is a balancing act around positive, neutral and negative buoyancy to maintain the best acceleration, momentum, velocity and drag conditions for each individual flight profile. However, possible optimum buoyancy at different points in ascending are described below:

1) Low to middle altitude - Generally positively buoyant so that you get "free" lift and that fuel is expended efficiently. Although to much "free" lift velocity at lower altitudes and you will lose out to air resistance proportionally to the extra amount of fuel that the "free" lift would have otherwise enabled to be carried.

2) Middle to high altitude - Generally when neutrally buoyant the fuel will be expended efficiently, but as an inverse of stage 1 in the preceding paragraph it may be sometimes more efficient to carry more fuel than to have "free" lift depending upon the most optimal drag, acceleration, momentum and velocity conditions needed at any point in time for each individual flight profile.

3) At high altitude - Generally it will be the best time to be negatively buoyant, as it is easiest and most optimal to negate most of the gravity lower down enabling the consequent ability to multiply fuel efficiency for the longest duration possible, if stages 1 and 2 in the preceding paragraphs were executed in the "best" way, the vehicle will be carrying a larger amount of secondary useful fuels, after obtaining a decent amount of momentum and velocity. It will then be able to expend that fuel in low drag environmental conditions.

The most efficient version for the flight profile of each craft is dependent upon the myriad of variables that will change within the flight as each vehicle ascends. This of course is dependent on empirical testing, but the advantage of using a craft 8, 98, 298, 308, 318 in accordance with the invention is that you can choose the most optimal efficient use of fuel for any point in time, enabling it we think to be far more efficient than any other system in current usage.

The speed of the crafts 8, 98, 298, 308, 318 is controllable, such control being related to the above described buoyancy control, optionally also by conventional means. Direction of the crafts is controllable by means described herein, optionally also by conventional means.

All orbital and suborbital trajectories will need the ability to control flight attitude to the appropriate angle for each flight. This can be done in a gross manner by changing LTACG flow to different sections of the engine, or to different engines within a multiple engine craft. Fine control can be done by conventional attitude control jets.

Moving lamina flow struts for Coanda and Liberator vehicles by the use of gimbals on either end of the struts, may also control flight attitude.

Ailerons on the lamina flow struts for the Coanda and Liberator vehicles, or on the air intake scoops of the Potent Voyager engine, may also control flight attitude. This is a selection of possible control mechanisms options that could be used individually or in combination.

The crafts 8, 98, 298, 308, 318 described herein can be provided with suitable steering means. Due to the flexibility of the crafts 8, 98, 298, 308, 318 of the invention, one could decide not to pressurise to 14 bar, but instead decide to carry more secondary fuels HTP and/or water enabling choosing more optimal flight profile characteristics, that are chosen for the cargo and/or passengers involved. For instance a rugged cargo would be able to take a high g acceleration saving fuel costs, but delicate glassware that could not currently be taken to orbit (due to the excessively violent vibrations of current launches, any delicate equipment would have to have excessive protective packaging to protect said equipment, making it practically and economically not viable, because of the increase in weight) could with our system be taken, using a more expensive fuel profile (though such a fuel profile is still likely to be very much cheaper than costs from alternative transport means available at the time of the invention). Again, trained healthy astronauts would likely use a higher g acceleration profile than elderly tourists again changing the costs involved. By using a combination of lighter than air fuels and heavier than air fuels we can select both optimal points on our fuel consumption and/or optimal in comfort and/or capability using the same basic equipment.

An example of a simple step by step descent profile would be:- 1) In orbit use some thrust to slow down from orbital velocity.

2) This would cause the vehicle to hit the outer atmosphere and enable gravity to reassert its vertical aspect in regards to the Earth.

3) Air resistance will slow the vehicle down.

4) The very large volume/surface area to mass ratio enables a slower descent at a shallower angle of descent, enabling the heat of re-entry to dissipate over a longer period of time.

5) This therefore enables a longer, slower, gentler descent to the ground, which can be either a passive or active landing, depending upon the amount of buoyancy the craft has at ground level. The large surface area of the wings or the crafts in general make them suitable to being covered with (thin film) PV solar cells, or Photovoltaic paint, or the like, which would enable the crafts to become self fuelling, by extraction of water and gases from the atmosphere and converting the constituents to the various optimal fuels previously described, thereby making the crafts cheaper to run.

Overall /sp equation proof follows. This proof is accurate for all embodiments of the various forms of craft (Potent Voyager, Coanda and Liberator) that contain a volume of LTACG.

/sp is traditionally calculated using the formula 7sp=Ve/gO, where :-

/sp is the specific impulse measured in seconds, ve is the average exhaust speed along the axis of the engine (in ft/s or m/s), gO is the acceleration at the Earth's surface (in ft/s2 or m/s2).

Whilst a launch craft in accordance with the invention is displacing more than its weight in air, then gO can be cancelled in the above equation giving the new equation 7sp=Ve + The amount of extra lift in m/s. This then will be defined as /spl , which is used in the Overall /sp equation below. Whilst a launch craft in accordance with the invention is displacing exactly its own weight in air, then gO can be cancelled in the above equation giving the new equation 7sp = Ve. This then will be defined as 7sp2, which is used in the Overall Isp equation below.

Whilst a launch craft in accordance with the invention is displacing a proportion of its own weight in air, then gO can be cancelled for a proportion of its weight at that specific point in time in units of seconds averaged out with the proportion of weight that is not displaced using the original Isp formula giving a value 7sp3 calculated using the new equation:-

( 7sp2 * MassKgOfAirDisplaced ) + ( Isp * ( GmolvatapsKg - massKgofairdisplaced ))

7sp3 =

GmolvatapsKg

Where GmolvatapsKg = Gross mass of launch craft in accordance with the invention at a particular second in time. (Note this includes weight of fuel, which can change on a per second basis).

7sp3 is used in the Overall Isp equation below.

Whilst the vehicle is out of the atmosphere and is still accelerating towards orbital velocity. The traditional way of calculating Isp needs to be used. Therefore hence the expression Isp=ve/g0 is used in the Overall Isp equation below. So the Overall Isp for a launch craft in accordance with the invention in a simplified form can be expressed as:-

(/spl * tl) + ( 7sp2 * t2 ) + ( 7sp3 * t3 ) + ( 7sp * t4 )

Overall 7sp =

t=time of flight in seconds = tl + 12 + 13 + 14

If a craft is initially heavier than air, and the initially heavier than air craft is continuously heavier than air throughout its flight profile, although the craft still contains a volume of LTACG, the 7spl and /sp2 components of the Overall /sp equation can be removed. If at any stage the craft becomes lighter than air, all parts of the Overall /sp equation are needed.

Note /spl and /sp3 will vary on a per second basis, due to lift from displacement being a variable on any given second. So the overall /sp is a simplified equation, and the traditional proportion of /sp will have minor changes due to gO changing slightly with changes in height until 1st orbital height has been reached.

The above proof shows that the theoretical maximum /sp of burning H₂ in air when the launch vehicle is displacing its own weight in air should be approximately 80,000 seconds at around Mach 5 to 6 and under the same conditions.

Waste heat dumped into H₂0 gives an approximate /sp of 1900 seconds. H₂0₂ by itself gives an approximate /sp of 1610 seconds at 1 atmospheric pressure. H₂0₂ used with H₂ gives an approximate /sp of 3519 seconds at 1 atmospheric pressure.

Summary The above are general basic versions for getting mass to L.E.O. and above cheaply. Like the internal combustion engine that enabled cars, vans, trucks, tanks, and aeroplanes, the concept of the invention can take many forms not limited to the primary three versions outlined (Potent Voyager, Coanda, and Liberator class). The basic concept of the invention is that the primary fuel lifts the whole system and initiates positive feedback loops, by taking advantage of the differences within the environment that it is designed initially to ascend in.

By using: a) Archimedes displacement. b) Pressure differentials. c) Utilising the usual disadvantage of frictional heat to increase the energy state of the primary fuel, thereby reducing efficiency losses. d) Both primary and any secondary fuels are utilised more efficiently because substantially all acceleration provided by them go into accelerating the vehicle and payload, without the need to expend any of their energy in holding up the vehicle/payload and fuel against gravity while the vehicle is lighter than air mass displaced. That contrasts with a continuous energy loss on a per second basis over the time it takes to reach orbital velocity in traditional rocketry. This also means that the engine size can be relatively small compared to the size and mass of the vehicle, giving a virtuous circle of efficiencies, because the engine has less work to do at any one point of time, given the nature of the advantages that are noted above. e) Burn rate and therefore thrust can be optimised to the changing environment and weight conditions of the vehicle during ascent. The vehicle can also ascend at any point without burning fuel, whilst it is lighter than air. Take off would be a natural point to do so, on the grounds of safety, but other points could also be utilised to customise and optimise the ascent profile. f) An addition to the craft concept of the invention buoyant and neutrally buoyant modes of operation is that, the craft concept of the invention system of using lighter than air combustible and/or non-combustible gases both actively and passively, can also be used in initially heavier than air launch vehicles, and also in continuously heavier than air launch vehicles, which would improve the power to weight ratio compared to traditional launch vehicles.

Initially heavier than air launch vehicles with integrated craft concept of the invention systems can take advantages of some of the efficiency of the craft concept of the invention system. This additional hybridisation of the technologies involved is advantageous in that it gives greater options in the configurations of weight and volume without major changes to standard craft concept of the invention launch vehicles, by giving greater options for the configurations of payload mass, and flight profile trajectories.

There are a number of safety features inherently in the craft concept of the invention system:-

1) There will be no problems with cryogenic boil off.

2) If launched passively the concept of the invention gives plenty of time for problem solving/escape for manned missions and recovery of payload for all mission types. 3) The concept of the invention can have a solid state/liquid safety valve for backblast protection (see figs. 11 & 12). The Tesla™ solid state valve can be used as replacement for, or in conjunction with, our safety valve system, and at the ends of our air scoops, to further reduce blowback losses.

4) The crafts of the invention are inherently stable whilst they are positively buoyant.

The holistic nature of all these various systems gives a great deal of flexibility to achieve various optimal sweet points for different launch and payload profiles, with customisable variants for achieving: Suborbital, Low Earth Orbit (L.E.O.), Polar orbital, Geostationary Earth Orbit (GEO) and Earth escape velocities.

Reference Information:

Claims

1. A flying craft comprising,

a lighter than air combustible gas receptacle,

the flying craft being made so that a lighter than air combustible gas within the lighter than air combustible gas receptacle can at least contribute to lifting the flying craft by the buoyancy of the lighter than air combustible gas and utilising Archimedes displacement principle,

the flying craft comprising a combustion engine or a combustion propulsion means, the combustion engine or the combustion propulsion means comprises a reaction chamber and a combustion chamber,

the lighter than air combustible gas receptacle being designed to be in gaseous communication with the combustion engine or combustion propulsion means, via ducting,

the flying craft being made so that, as the altitude of the flying craft increases, a pressure differential between a lighter than air combustible gas inside the a lighter than air combustible gas receptacle and gas or gases surrounding the outside of the craft, which can come into existence, or increase, as the altitude of the flying craft increases, is made to cause a lighter than air combustible gas to be forced out of the receptacle, via the reaction chamber, towards the combustion chamber of the combustion engine or combustion propulsion means,

the flying craft comprising an Oxygen receptacle or an Oxygen collector, the Oxygen receptacle or the Oxygen collector being designed to be in gaseous communication with the combustion engine or combustion propulsion means, via ducting,

the reaction chamber being designed to permit a chemical reaction between a lighter than air combustible gas from the lighter than air combustible gas receptacle and Oxygen from the Oxygen receptacle or the Oxygen collector, and the flying craft being arranged that such a chemical reaction provide thrust or lift to the flying craft,

the combustion chamber being designed to combust a lighter than air combustible gas from the lighter than air combustible gas receptacle in the manner of a fuel, and Oxygen from the Oxygen receptacle or the Oxygen collector, and to discharge a product or products of the combustion of a lighter than air combustible gas fuel and Oxygen out of an exhaust of the combustion chamber, so as to provide thrust or lift to the flying craft, and wherein the craft is made so that at least at one point during its flight both lighter than air combustible gas buoyancy, and thrust from combustion of lighter than air combustible gas, from the same source, contribute to lifting the craft.

2. A flying craft according to Claim 1, wherein the flying craft comprises an Oxygen collector, the Oxygen collector comprises one or more air deflector(s) arranged to channel air from the atmosphere into the combustion chamber of the combustion engine or combustion propulsion means.

3. A flying craft according to Claim 2, wherein air from the air deflector(s) is/are arranged to enter the combustion chamber of the combustion engine or combustion propulsion means at a different angle to the LTACG to allow turbulent mixing of air and LTACG when the air and LTACG meet.

4. A flying craft according to any one or more of the preceding claims, wherein the flying craft comprises a chamber for rocket grade Hydrogen peroxide (H₂0₂).

5. A flying craft according to Claim 4, wherein the flying craft comprises a catalyst to cause the rocket grade Hydrogen peroxide (H₂0₂) in the chamber to form 0₂, wherein the craft is designed so that 0₂ from the catalytic reaction, and LTACG can react, so as to provide steam, which can be discharged for a further form of thrust or lift to the flying craft.

6. A flying craft according to any one or more preceding claim, when dependent on claim 4, wherein the craft is designed so that H₂0₂ in the chamber can be used to absorb at least some of the heat from the H₂ + 0₂ (from air) reaction.

7. A flying craft according to any one or more preceding claim, when dependent on claim 4, wherein the craft is designed so that H₂0₂ in the chamber can be used to absorb at least some of the heat from the H₂ + 0₂ (from H₂0₂) reaction.

8. A flying craft according to claim 6 or 7, wherein the craft is designed so that Oxygen, a by product of using H₂0₂ as a coolant, can be used to react with LTACG to obtain a further source of thrust.

9. A flying craft according to any one or more of claims 4 to 8, when dependent on claim 4, wherein the craft is designed so that H₂0₂ in the chamber can be used as ballast.

10. A flying craft according to any one or more preceding claim, wherein the flying craft comprises a chamber for liquid coolant, such as water coolant, to absorb at least some of the heat from the H₂ + 0₂ (from air) reaction.

11. A flying craft according to any one or more preceding claim, when dependent on claim 4, wherein the flying craft comprises a chamber for liquid coolant, such as water coolant, to absorb at least some of the heat from the H₂ + 0₂ (from H₂0₂) reaction.

12. A flying craft according to claim 10 or 1 1 , wherein the craft is designed so that a coolant in the coolant chamber can be used as ballast.

13. A flying craft according to any one or more of claims 10, 11 and 12, wherein the flying craft is designed to discharge steam from the chamber for liquid coolant, for a further form of thrust.

14. A flying craft according to any one or more of claims 10 to 13, wherein the craft is designed so that coolant from the chamber for liquid coolant preheats the LTACG so the LTACG combusts more efficiently (because warmer reactants react more efficiently than cooler ones).

15. A flying craft according to any one or more preceding claim, wherein the craft comprises an inert gas receptacle which surrounds at least part, the majority, or substantially all of the LTACG receptacle, preferably the receptacle may have an inert gas therein, most preferably the inert gas is Helium.

16. A flying craft according to any one or more preceding claim, wherein the flying craft is arranged so that the LTACG receptacle is above the engine or combustion propulsion means during flight.

17. A flying craft according to any one or more preceding claim, wherein the flying craft comprises a receptacle designed to house a second, not lighter than air, combustible gas therein, in gaseous communication with the combustion engine or combustion propulsion means, so that the second combustible gas can be combusted.

18. A flying craft according to any one or more preceding claim, wherein the flying craft comprises a ring shaped cross section balloon, and an engine or combustion propulsion means mounted in the opening of the ring shaped cross section balloon.

19. A flying craft according to claim 18, wherein the surface on the inside of the ring shaped cross section balloon enlarges in diameter to take advantage of the Coanda effect.

20. A flying craft according to claim 18 or 19, wherein the craft comprises an engine or combustion propulsion means arranged inbound from the surface on the inside of the ring shaped cross section balloon, and spaced therefrom, the craft also comprises one or more vanes which direct air from the inside surface of the ring shaped cross section balloon inwardly towards the engine or combustion propulsion means.

21. A flying craft substantially as described herein and/or with reference to any one or more of Figures 1 to 5 of the drawings.

22. A flying craft substantially as described herein and/or with reference to any one or more of Figures 6 to 10, and 16 to 18, of the drawings.

23. A flying craft comprising a plurality of crafts in accordance with any one or more of the preceding claims joined together to make a larger craft.

24. A flying craft according to Claim 23, wherein the craft is designed to take off horizontally, increasing the pitch of the craft as it ascends, so the pitch is more vertical than horizontal.

25. A flying craft according to Claim 24, wherein once the craft has substantially passed the atmosphere, the pitch of the craft decreases, so the craft can make a horizontal/tangential orbit.

26. A flying craft substantially as described herein and/or with reference to any one or more of Figures 18 to 21 of the drawings.

27. A flying craft, or part thereof, substantially as described herein and/or with reference to any one or more of the drawings.

28. A method of operating a flying craft, the method comprising,

providing a craft according to any one or more of claims 1 to 25,

putting LTACG in the craft,

directing LTACG from the LTACG receptacle towards the combustion chamber of the combustion engine or combustion propulsion means, and

reacting LTACG and Oxygen to provide steam which is discharged so as to provide thrust or lift to the flying craft.