Satellite Launcher And Method For Putting Satellites Into Orbit Using Said Satellite Launcher

Abstract

The satellite launcher comprises a plurality of stages detachable from each other, at least one stage including at least one engine, and at least one of said stages carrying a payload, and said stages are placed one beside or around the other, so that the width of the vehicle is at least one third of its length.

The method comprises the following phases: a) ascent of the vehicle with a balloon from a ship; and b) ignition of engines of the vehicle to put a satellite placed in the vehicle into orbit.

Claims

1. A satellite launcher comprising a plurality of stages detachable from each other, at least one of said stages including at least one engine, and at least one of said stages carrying a payload, wherein said stages are placed either beside or around one another, so that the width of the vehicle is at least one third of its length.

2. The satellite launcher according to claim 1, wherein the width of the vehicle is the same or greater than the length of the vehicle.

3. The satellite launcher according to claim 1, wherein at least one of the stages comprises at least one tank.

4. The satellite launcher according to claim 1, wherein one of said stages is a central stage which is surrounded by one or more additional torus-shaped stages.

5. The satellite launcher according to claim 3, wherein each tank is a torus-shaped tank.

6. The satellite launcher according to claim 4, wherein each additional stage comprises a plurality of engines equidistantly spaced apart defining a circle.

7. The satellite launcher according to claim 1, wherein the payload is placed in the central stage.

8. The satellite launcher according to claim 6, wherein the payload is protected by a fairing which could be is either detachable or retractably attached to any of the stages.

9. The satellite launcher according to claim 4, wherein the central stage also comprises a guidance, navigation and control system.

10. The satellite launcher according to claim 3, wherein each tank is made of composite.

11. The satellite launcher according to claim 3, wherein the tank of at least one stage is connected to at least one engine of another stage.

12. The satellite launcher according to claim 4, wherein the tank of the most external stage is connected to the engines of this stage and to the engines of the rest of stages.

13. A method for putting satellites into orbit using the satellite launcher according to claim 1, wherein the method comprises the following phases: a) ascent of the vehicle with a balloon from a ship; and b) ignition of engines of the vehicle to put a satellite placed in the vehicle into orbit.

14. The method according to claim 13, wherein the ascent of the balloon with the vehicle takes from 80 and 100 minutes and places the vehicle at a height from 15 to 25 km (50,000 to 83,000 ft).

15. The method according to claim 13, wherein the ignition of the engines of the vehicles comprises at least the following steps: a first step to place the vehicle at a height of about 80 km (263,000 ft or 50 mi), detaching a first stage of the vehicle; a second step to place the vehicle at a height of about 300 km (1,000,000 ft or 187 mi), detaching a second stage of the vehicle; and a third step to or the satellite at a height of 600 km (2,000,000 ft or 373 mi) performing several fires and detaching the central stage from the satellite.

16. The method according to claim 15, wherein the first step lasts about 120 seconds and it takes the vehicle at an inertial speed of about 2.8 km/s (6,264 mph).

17. The method according to claim 15, wherein the second step lasts about 150 seconds and it takes the vehicle at an inertial speed of about 5 km/s (11,185 mph).

18. The method according to claim 15, wherein the third step comprises a first fire that lasts from 90 and 110 seconds, and second fire that lasts from 180 and 220 seconds and a final fire that lasts from 140 and 160 seconds.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0048] For a better understanding of what has been disclosed, some drawings are attached, showing diagrammatically and only as a non-limitative example, one embodiment of the invention.

[0049] FIG. 1 is an exploded perspective view of the satellite launcher according to one embodiment of the invention;

[0050] FIG. 2 is an elevation view of the satellite launcher according to the preferred embodiment the invention;

[0051] FIG. 3 is an elevation view of a system for putting satellites into orbit used for the method according to the present invention;

[0052] FIG. 4 is a diagrammatical view of the flight train used in the system for putting satellites into orbit;

[0053] FIG. 5 is a diagrammatical view of the connections between the tanks and the engines of the aerospace according to the invention in three different phases during the method according to the invention; and

[0054] FIG. 6 is a diagrammatical view of the whole method for putting satellites into orbit according to the present invention.

DESCRIPTION OF A PREFERRED EMBODIMENT

[0055] The present invention relates to a satellite launcher, specifically a suborbital and orbital launcher, and to a method for putting satellites into orbit using said vehicle.

[0056] The satellite launcher 10 according to the present invention comprises several stages which make up a blunt ensemble and are themselves blunt. In the preferred embodiment there are two torus-shaped stages and a central stage. These stages are detachable from each other during the method according to the present invention, as described hereinafter. These stages are preferably coated with adequate re-entry material, which could be ablative, such as phenolic resins or radiative such as carbon-carbon composites or carbon aerogels.

[0057] According to the embodiment of FIG. 1, a first stage 1 comprises a torus-shaped structural tank 11 and wraps a second stage 2, which also comprises a torus-shaped tank 21, and a third or central stage 3, which has an ellipsoidal tank 31. This configuration saves dry mass since all the engines can be ignited and contributing to the thrust during the whole trajectory.

[0058] Furthermore, the vehicle according to the invention uses composite tanks 11, 21, 31 that reduce weight, costs, from using conventional metal tanks.

[0059] All three stages 1, 2, 3 comprise corresponding engines 12, 22, 32, distributed in a ring-like symmetrical pattern in the first and second stages 1, 2, and a core engine 32 in the third stage 3. The propellants combination chosen to be used in the preferred embodiment is liquid oxygen and liquid methane. This bi-propellant combination is the perfect match between performance, simplicity of hydrocarbon combustion and green propulsion. This patent should cover other propellants, both monopropellant and bi-propellants.

[0060] Preferably, the engines 12, 22, 32 are pressurized by a suitable pressurant gas, such as Helium, either compressed, liquefied or enriched with reactants (as in the Tridyne method disclosed in U.S. Pat. No. 3,779,009 A). In another instance, the engines 12, 22, 32 are pressurized by a Vapor Pressurization (VaPak) system. The main advantage of the VaPak concept compared to alternative like inert gas pressurization or pumped-fed systems is its reduction in complexity and a reduction in empty weight. By using the VaPak system the vehicle according to the invention is expected to have a high performance while remaining a simple system.

[0061] Vapor Pressurization is based on the use of high vapor pressure of the propellants to provide the pressure difference required for the propellants to flow into the combustion chamber. The high vapor pressure is obtained from the internal energy of a liquid stored in a closed container. As propellant is drained, liquid boils, and resulting gas re-pressurizes the propellant tank. The pressurization system is used at every stage to pressurize the propellants tanks to a pressure large enough for the propellants to flow into the combustion chambers of the rocket engines.

[0062] Using a propulsion system based on green propellants and a first balloon ascent, as will be described hereinafter, with inert Helium, will mitigate the chemical contamination of traditional launchers due to the presence of ozone depleting chemicals on the rocket combustion products, which affect the ozone layer that protects all life forms from solar UV radiation.

[0063] The satellite launcher according to the present invention also comprises a GNC (Guidance, Navigation and Control) system 33 at least in the central stage 3, in which it is also placed a payload 4, protected by a corresponding fairing 41. This payload 4 includes the satellite. There is also a standard adaptor for the satellite to be attached and released from the central stage.

[0064] The compact configuration of the vehicle according to the invention makes control simpler than the one of traditional very slender bodies. The GNC system 33 calculates the optimum trajectory and controls the elements for modifying the trajectory. The software implemented would be capable of controlling the vehicle in the event of engine failure (the mission success is assured even in the event of a one-engine-out of the first stage and/or second stage), as will be explained hereinafter.

[0065] The method for putting satellites according to the invention comprises two different phases: balloon ascent and vehicle ignition, as shown in FIGS. 3 and 6.

[0066] The satellite launcher 10 according to the invention is preferably launched from a ship 6 reducing the risk of launch delays by avoiding bad weather and also compensating ground winds and adapting the engine ignition spot to mission and safety requirements.

[0067] The balloon 5 during the first phase of the flight cycle carries the vehicle according to the invention up to 20 km and the ascent lasts around 90 minutes. The balloon 5 will be filled with a suitable lifting gas through corresponding inflation tubes 7 such as Helium or Hydrogen. In one instance a hot air balloon may be used instead of a gas balloon.

[0068] The buoyancy effect takes the balloon 5 with the vehicle 10 to a predefined altitude between 20 km and 25 km. In case of a malfunction of the rocket or satellite prior to its detachment from the balloon, the whole mission could be aborted with recovery, since the balloon may controllably vent gas from its apex valve and descend to the sea where the payload and rocket may be re recovered for inspection or future re-flight.

[0069] The second phase of the flight cycle starts once the engines of the vehicle 10 are ignited, and it consists of several stage firings different steps (the exact values may change depending on the orbital destination of the flight, and this is just an illustrative case):

[0070] The first step lasts 120 seconds and it takes the vehicle 10 from 20 to 80 km (66,000 to 263,000 ft) at an inertial speed of 2.8 km/s (6,264 mph). During this step the engines of the vehicle 10 are producing thrust with a total vacuum impulse of 104 kN (23,380 lbf). The cover 41 that protects the payload 4 is detached, or retracted at about the same time that the first stage 1 separates from the rest of the vehicle.

[0071] The second step raises the vehicle 10 to 300 km (33,000 ft to 187 mi) in 150 seconds and the second stage 2 is detached, and at the end of this step the vehicle 10 is flying at an inertial speed of 5.1 km/s (11,500 mph). During the second step the engines of the vehicle 10 produce thrust with a maximum vacuum impulse of 14 kN (3,148 lbf).

[0072] The last step performs several fires to optimally orbit the payload 4. The first fire lasts for 100 seconds and allows the payload 4 to reach 600 km (373 mi) of altitude while still slightly below the target orbital speed. Then, the third stage 3 coasts for 200 seconds and a final fire of 145 seconds optimally orbits the payload 4. Lastly, the last boost of the third stage 3 is performed to detach and de-orbit the third stage 3 in order to minimize the amount of space debris left by the mission. This method is shown in FIG. 6.

[0073] The stages 1, 2, 3 may brake up in re-entry or may be recovered by having them either land on land or on a barge in the sea as originally described on the publication by Yoshiyuki Ishijima et al., Re-entry and Terminal Guidance for Vertical-Landing TSTO (Two-Stage to Orbit), AAIA Pub. No. 98-4120 in 1998. In both cases, a large net may be used to simplify the guidance requirements and have the stage just fall into the net, and not land as precisely as it would be required on a flat Helipad sort of surface. Using the net also saves on dry weight of the stages, since they would not require landing legs.

[0074] The main advantage of using a boat to launch a balloon is to balance the wind speed with the ship speed, so there is zero relative wind speed between the air and the balloon, which simplifies the operation. This also provides the flexibility to launch from most of the surface of the planet, which is covered by water, better meeting mission needs than launching from a fixed spaceport.

[0075] The ship, by moving at the same speed as the wind, creates a near zero wind column for inflating and releasing the balloon from the deck. The ship itself does not need any significant adaptation for the operation and any ship with a sufficiently big flat area to accommodate the bubble of the balloon being inflated, and with the right conditions for propellant storage, could be rented to perform the flight. The payload is mounted near the balloon inflation area.

[0076] The balloon 5 also comprises a flight train and gondola that remain attached to the balloon and includes its own avionics system, shown in FIG. 4, which is responsible for sensing the motion of the vehicle, monitoring the subsystems state, communicating with ground and providing the computational power and internal communications for the GNC system 33 and subsystems interactions. The avionics system also supplies the necessary electrical power satisfying the need for high power when required and enabling ground to on-board switching. Communication between ground and the vehicle 10 is crucial during the mission; vehicle-to-ground communications provide data to the ground concerning real-time flight data towards vehicle on ground tracking and monitoring, and off-line data for post-mission exploitation; ground-to-launcher communications provide link to ground safety commands.

[0077] The flight train is located between the balloon 5 and the vehicle 10 and hosts all the necessary equipment for a successful balloon operation comprising at least the following elements: [0078] a GPS 51 for detecting the position of the balloon 5; [0079] one or more transponders 52 for coordination with Air Traffic Control; [0080] one or more telemetry systems 53; [0081] one or more radar reflectors 54 to comply with the Rules of the Air; [0082] a flight termination system 55 to ensure separation of the payload; [0083] optionally parachute 56 including a parachute releasing system 57 to recover the gondola and flight train in case of balloon failure; [0084] a ballast machine 58 to precisely control the altitude; [0085] a truck plate 59 to transmit the loads; and [0086] a mechanical adaptor 42 to engage with the vehicle below 4.

[0087] As shown in FIG. 5, in order to optimize the capabilities of the launcher, a cross-tanking capability would be implemented in some cases, so that the remaining stages are full when a stage separates, in the preferred embodiment this means having piping from the 1st to the 2nd and 3rd and from the 2nd to the 3rd. This piping disconnects (possibly by a normally open pyrovalve) when the separation takes place.

[0088] During the first step of the second phase of the method according to the invention, the engines 12, 22, 32 of the first, second and third stages 1, 2, 3 are fed by the tank 11 of the first stage 1.

[0089] During the second step of the second phase of the method according to the invention, the first stage 1 has been detached and the engines 22, 32 of the second and third stages 2, 3 are fed by the tank 21 of the second stage 2.

[0090] During the third step of the second phase of the method according to the invention, the second stage 2 has been detached and the engine 32 of the third stages 3 is fed by the tank 31 of the third stage 2.

[0091] A breakdown of the masses of each stage of the vehicle according to the invention is shown in Table I. The table also illustrates the ideal increment of velocity (deltaV) that every stage contributes to, this is an example, numbers may vary.

TABLE-US-00001 TABLE I Masses breakdown First stage Second stage Third stage Structural mass (kg) 552.7 118.8 103.7 (lbs) 1,219 261.9 228.6 Fairing (kg) 25 0 0 (lbs) 55 0 0 Propellant mass (kg) 3284.6 622.6 218.7 (lbs) 7,241.3 1,373 482.2 Total stage mass (kg) 3862.3 741.4 322.4 (lbs) 8,514.9 1,635 710.8 Stage Payload (kg) 1138.8 397.4 75 (lbs) 2,510.6 876.1 165 Engine Isp (s) 342 342 342 Ideal deltaV (m/s) 3587.8 2654.6 2681.4 deltaV contrib. (%) 40.2 29.8 30

[0092] Even though reference has been made to a specific embodiment of the invention, it is clear for a person skilled in the art that the disclosed system is susceptible of a number of variations and modifications, and that all the details mentioned can be substituted by other technically equivalent ones, without departing from the scope of protection defined by the attached claims.