WO2023213420A1

WO2023213420A1 - A propulsion system for a spacecraft and method for pressure feeding

Info

Publication number: WO2023213420A1
Application number: PCT/EP2022/068647
Authority: WO
Inventors: Felix SCHILY
Original assignee: Deltaorbit Gmbh
Priority date: 2022-05-02
Filing date: 2022-07-05
Publication date: 2023-11-09

Abstract

The invention relates to a, preferably pumpless, pressure fed propulsion system for a spacecraft, preferably a satellite or microsatellite, the system comprising: - a propellant tank (101) for storing a propellant; - a first heat exchanger (105) that is fluidly connected to the propellant tank and arranged downstream thereof so as to be able to heat the propellant with waste heat from the thruster; - a second heat exchanger (107) that is fluidly connected to the first heat exchanger and arranged downstream thereof so as to be able to heat the stored propellant; and - a thruster (01) that is fluidly connected to the second heat exchanger and arranged downstream thereof so as to be able to generate thrust using the propellant.

Description

A PROPULSION SYSTEM FOR A SPACECRAFT AND METHOD FOR PRESSURE FEEDING

FIELD OF THE INVENTION

The invention relates to a propulsion system for a spacecraft and a method for pressure feeding.

BACKGROUND

Chemical in-space propulsion systems have used highly toxic and carcinogenic propellants for more than fifty years. In these systems, hydrazine is used - either as monopropellant or as bipropellant with a suitable oxidizer (often nitrogentetroxide; short: NTO). These propellants are usually stored as sub-cooled liquids in tanks with negligible vapor-pressure. In large upper-stage engines often pumps are used for transporting these propellants from tanks to the combustion chamber.

Another system is a separate inert gas stored in dedicated high-pressure tank that is used to drive the fluids from propellant tanks to engine. While these systems have been used very successfully throughout the decades, the high toxicity of the propellants has increasingly become a burden.

In recent years the development of “green” propellant alternatives has gained considerable traction, driven partly by the emerging small- and microsatellite industry which is innovating to avoid the extreme safety measures necessary to handle hazardous propellants like hydrazine.

Additionally, the propulsion requirements for these types of satellites are usually small, which means that these systems can make do with relatively inefficient engines. While a new generation of green propulsion systems based on N2O, H2O2, ammonium dinitramide (ADN) or hydroxylammonium nitrate (HAN) can serve these particular use cases, the respective technology currently does not scale well to larger engines.

In order to reach higher thrust levels, technology derived from those more commonly found in large launch vehicles can be adapted and scaled down for use in satellites. The cryogenic propellants methane and oxygen for example can be stored indefinitely under space-conditions and exhibit very high efficiency.

In contrast to classical, storable propellants like hydrazine/NTO these are stored at much lower temperatures and exhibit considerably vapor-pressure at medium cryogenic temperatures. These self-pressurization properties can be utilized to derive new methods for driving the propellants from tanks to engine.

To enable a propellant to enter a thrust chamber, the propellant must have a greater pressure than the pressure in the thrust chamber. This can be achieved by raising the pressure by means of pumps (“pump-fed”), by storing the propellant at a sufficiently high pressure (“pressure-fed”), or by a combination thereof.

Most commonly, large lower-stage engines are pump-fed, whereas smaller engines used for maneuvering, on upper stages, or on satellites are pressure-fed. They have in common that the propellants are delivered to the engine in liquid form. Pump-fed engines suffer from high development and manufacturing cost due to the complexity of turbo pumps, whereas the simpler pressure-fed systems have an increased mass, since they require an additional high-pressure tank to store the pressurization gas. The pressurization gas must have a significantly higher pressure than the propellants, since after expanding into the propellant tank its pressure must still be sufficient to ensure the propellant supply to the thrust chamber at design pressure.

There are various examples of systems where the designers circumvented the choice between the system complexity of turbo pumps and the added mass of a high-pressure tank for the pressurization gas by evaporating some of the propellant in its tank in order to maintain pressure. These systems are called “selfpressurized”.

There are systems that withdraw propellants as saturated liquid from a tank in two-phase equilibrium. The subsequent pressure loss causes boiling in the tank, which produces new gaseous fuel and thus partially mitigates the pressure loss. This method allows to almost empty the propellant tank while merely losing about 30% of the original pressure.

However, as pressure loss in the feed lines further downstream occurs, more liquid is evaporated and bubbles occur, leading to challenging conditions regarding reproducibility and combustion stability. If the propellants are withdrawn in gaseous form, much more heat is required to evaporate the propellants. Without an additional heat source, the pressure in the propellant tanks usually drops too fast for a realistic use of this principle.

On the Space Shuttle, for example, both propellants in the external tank were pressurized by adding gaseous propellant of the same kind, that had previously been heated using the pre-burner exhaust. On the TITAN 34D gas generators were used to pressurize the propellant tanks. Both these systems still require powerful turbo pumps that allow to feed propellants - once withdrawn - back into the tanks.

US 3 570 249 A discloses similar type of pressurization - feeding some exhaust gases into the propellant tanks, which in addition uses jet pumps driven by monergol propellants. The tank pressurization merely aims at avoiding cavitation in the pumps further downstream.

US 7 784 269 B1 discloses a system that uses a coolant from a separate coolant tank to cool the engine and subsequently transport the heat to the propellant tank, where it is used to evaporate enough propellant such that the tank pressure assumes the desired value. It is suggested to either route the coolant in a closed loop requiring a pump, to vent the coolant overboard, or to add the coolant - which may be combustible or not - to the combustion chamber to increase thrust, which requires a pump or sufficient coolant tank pressure. Combinations of the three options allow to adjust the supply of heat to the propellant tank such that the desired tank pressure is achieved.

Similarly, US 6 052 987 A discloses a coolant cycle which can also be used to pre-heat propellant. In all of these cases either a coolant pump is required, or another limited resource (the coolant) is introduced. Furthermore, an additional coolant tank is required and thus offers only limited advantages over traditional pressure-fed systems. US 2011 10 005 193 A1 discloses a pool boiling heat exchanger in order to simplify combustor design. Whereas also here, waste heat is used to pre-heat the propellant(s), this configuration does not prevent pressure loss in the propellant tank and requires a separate system for tank pressurization.

US 9 446 862 B2 discloses using the waste heat of main and auxiliary thrust chambers to maintain and control the pressure in the propellant tanks. This is accomplished by cooling the respective combustion chamber with liquid propellant, which is evaporated in the process and then routed back to the propellant tank using a pump. While this system may reduce the necessary pump power, this advantage does not offset its additional complexity.

SUMMARY OF THE INVENTION

It is the object of the invention to improve feed systems for (chemical) rocket engines. The object is achieved by the subject-matter of the independent claims. Preferred embodiments are subject-matter of the dependent claims.

The invention provides a, preferably pumpless, pressure fed propulsion system for a spacecraft, preferably a satellite or microsatellite, the system comprising: a propellant tank for storing a propellant; a first heat exchanger that is fluidly connected to the propellant tank and arranged downstream thereof so as to be able to heat the propellant with waste heat from the thruster; a second heat exchanger that is fluidly connected to the first heat exchanger and arranged downstream thereof so as to be able to heat the stored propellant; and a thruster that is fluidly connected to the second heat exchanger and arranged downstream thereof so as to be able to generate thrust using the propellant.

Preferably, the propellant tank comprises a gaseous outlet that is arranged such that when the thruster is active only gaseous propellant is discharged from the gaseous outlet. Preferably, the system further comprises a valve assembly that is configured to control a gaseous propellant flow from the first heat exchanger to the second heat exchanger and/or to the thruster.

Preferably, the valve assembly comprises a gaseous bypass valve that is arranged to control a direct gaseous propellant flow from the first heat exchanger to the thruster. Preferably, the valve assembly comprises a heat exchanger control valve that is arranged to control a direct gaseous propellant flow from the first heat exchanger to the second heat exchanger. Preferably, the valve assembly comprises a splitting valve that is arranged to receive a gaseous propellant flow directly from the first heat exchanger and controllably discharge a split gaseous propellant flow directly to the second heat exchanger and to the thruster. Preferably, the valve assembly comprises a mixing valve that is arranged to receive a gaseous propellant flow directly from the first and second heat exchangers and controllably discharge a gaseous propellant flow mixed therefrom to the thruster.

Preferably, the propellant tank includes a liquid outlet that is arranged such that when the thruster is active only liquid propellant is discharged from the liquid outlet. Preferably, the system further comprises a controllable liquid bypass that fluidly connects the liquid outlet to the thruster so as to feed liquid propellant to the thruster in a controllable manner.

Preferably, the system further comprises an auxiliary thruster that is fluidly connected to the propellant tank, preferably a gaseous flow from the propellant tank, in a controllable manner.

Preferably, the propellant tank is a primary propellant tank for storing a primary propellant. Preferably, the first heat exchanger is a primary branch thruster heat exchanger that is fluidly connected to the primary propellant tank and arranged downstream thereof so as to be able to heat the primary propellant with waste heat from the thruster. Preferably, second heat exchanger is a primary branch tank heat exchanger that is fluidly connected to the primary branch thruster heat exchanger and arranged downstream thereof so as to be able to heat the stored primary propellant.

Preferably, the system further comprises a secondary propellant tank for storing a secondary propellant. Preferably, the system further comprises a secondary branch thruster heat exchanger that is fluidly connected to the secondary propellant tank and arranged downstream thereof so as to be able to heat the secondary propellant with waste heat from the thruster. Preferably, the system further comprises a secondary branch tank heat exchanger that is fluidly connected to the secondary branch thruster heat exchanger and arranged downstream thereof so as to be able to heat the stored secondary propellant. Preferably, the thruster is fluidly connected to the primary and secondary branch tank heat exchangers and arranged downstream thereof so as to be able to generate thrust using the primary and secondary propellants.

Preferably, the primary propellant tank comprises a primary gaseous outlet that is arranged such that when the thruster is active only gaseous primary propellant is discharged from the primary gaseous outlet.

Preferably, the secondary propellant tank comprises a secondary gaseous outlet that is arranged such that when the thruster is active only gaseous secondary propellant is discharged from the secondary gaseous outlet.

Preferably, the valve assembly is a primary valve assembly that is configured to control a gaseous primary propellant flow from the primary branch thruster heat exchanger to the primary branch tank heat exchanger and/or to the thruster.

Preferably, the primary valve assembly comprises a primary gaseous bypass valve that is arranged to control a direct gaseous primary propellant flow from the primary branch thruster heat exchanger to the thruster. Preferably, the primary valve assembly comprises a primary heat exchanger control valve that is arranged to control a direct gaseous primary propellant flow from the primary branch thruster heat exchanger to the primary branch tank heat exchanger. Preferably, the primary valve assembly comprises a primary splitting valve that is arranged to receive a gaseous primary propellant flow directly from the primary branch thruster heat exchanger and controllably discharge a split gaseous primary propellant flow directly to the primary branch tank heat exchanger and to the thruster. Preferably, the primary valve assembly comprises a primary mixing valve that is arranged to receive a gaseous primary propellant flow directly from the primary first and second heat exchangers and controllably discharge a gaseous primary propellant flow mixed therefrom to the thruster. Preferably, the system further comprises a secondary valve assembly that is configured to control a gaseous secondary propellant flow from the secondary branch thruster heat exchanger to the secondary branch tank heat exchanger and/or to the thruster.

Preferably, the secondary valve assembly comprises a secondary gaseous bypass valve that is arranged to control a direct gaseous secondary propellant flow from the secondary branch thruster heat exchanger to the thruster. Preferably, the secondary valve assembly comprises a secondary heat exchanger control valve that is arranged to control a direct gaseous secondary propellant flow from the secondary branch thruster heat exchanger to the secondary branch tank heat exchanger. Preferably, the secondary valve assembly comprises a secondary splitting valve that is arranged to receive a gaseous secondary propellant flow directly from the secondary branch thruster heat exchanger and controllably discharge a split gaseous secondary propellant flow directly to the secondary branch tank heat exchanger and to the thruster. Preferably, the secondary valve assembly comprises a secondary mixing valve that is arranged to receive a gaseous secondary propellant flow directly from the secondary first and second heat exchangers and controllably discharge a gaseous secondary propellant flow mixed therefrom to the thruster.

Preferably, the liquid outlet is a primary liquid outlet that is arranged such that when the thruster is active only liquid primary propellant is discharged from the primary liquid outlet. Preferably, the liquid bypass is a primary controllable liquid bypass that fluidly connects the primary liquid outlet to the thruster so as to feed liquid primary propellant to the thruster in a controllable manner.

Preferably, the secondary propellant tank includes a secondary liquid outlet that is arranged such that when the thruster is active only liquid secondary propellant is discharged from the secondary liquid outlet. Preferably, the system further comprises a controllable secondary liquid bypass that fluidly connects the secondary liquid outlet to the thruster so as to feed liquid secondary propellant to the thruster in a controllable manner.

Preferably, the system further comprises a branch heat exchanger that is arranged to exchange heat between a gaseous primary propellant flow and a gaseous secondary propellant flow. Preferably, the branch heat exchanger is arranged in a primary propellant circuit between the primary branch tank heat exchanger and the thruster and in a secondary propellant circuit between the secondary propellant tank and the secondary branch tank heat exchanger, preferably a secondary valve arrangement.

The invention provides spacecraft, preferably a satellite or a microsatellie, comprising a preferred propulsion system.

The invention provides a method for pressure feeding a thruster of a propulsion system for a spacecraft, preferably a satellite or microsatellite, the method comprising: feeding gaseous propellant from a propellant tank to a first heat exchanger; with the first heat exchanger heating the propellant with waste heat from a thruster and feeding the heated propellant to a second heat exchanger; with the second heat exchanger heating the remaing propellant in the propellant tank and subsequently feeding the propellant from the second heat exchanger to the thruster, and generating thrust with the propellant.

Preferably, when the thruster is active in zero-g or microgravity, only gaseous propellant is discharged from the propellant tank through a gaseous outlet due to acceleration by the thrust.

Preferably, the method further comprises a valve assembly controlling a gaseous propellant flow from the first heat exchanger to the second heat exchanger and/or to the thruster.

Preferably, the valve assembly comprises a gaseous bypass valve that is controlling a direct gaseous propellant flow from the first heat exchanger to the thruster. Preferably, the valve assembly comprises a heat exchanger control valve that is controlling a direct gaseous propellant flow from the first heat exchanger to the second heat exchanger. Preferably, the valve assembly comprises a splitting valve that is receiving a gaseous propellant flow directly from the first heat exchanger and is controlling discharging of a split gaseous propellant flow directly to the second heat exchanger and to the thruster. Preferably, the valve assembly comprises a mixing valve that is receiving a gaseous propellant flow directly from the first and second heat exchangers and is controlling discharging of a gaseous propellant flow that is being mixed therefrom to the thruster.

Preferably, when the thruster is active in zero-g or microgravity, only liquid propellant is discharged from the propellant tank through a liquid outlet due to acceleration by the thrust. Preferably, the system further comprises a controllable liquid bypass that feeds liquid propellant to the thruster in a controllable manner.

Preferably, an auxiliary thruster is controllably fed from the propellant tank, preferably with a gaseous flow from the propellant tank.

Preferably, branch heat exchanger exchanges heat between a gaseous primary propellant flow and a gaseous secondary propellant flow.

The present disclosure generally relates to thrusters, such as rocket engines, more specifically to the process and apparatus ensuring the cooling and fuel supply of the thruster using at least one propellant that is stored as a liquid or as a mixture of liquid and gaseous phase.

With this new measures for transporting gaseous or liquid propellants from storage tanks into a combustor, neither pumps nor a high-pressure storage tank are required. It is most suitable for in-space chemical propulsion systems, where combustor pressure is comparatively low.

In contrast to classical feed systems the described system stores one or more propellants in a self-pressurizing two-phase state. By controlling the temperature of the stored propellants, the tank pressure can be adjusted. To reduce the probability for boiling and/or for two-phase flow in feed lines, valves, and injectors - which can have negative effects on system reliability - the propellants are preferably taken from the gaseous phase in the tanks.

The gaseous propellants then flow through a heat exchanger around the combustion chamber, where they can pick up waste heat. The warm propellants are then preferably routed to a second heat exchanger. The second heat exchanger is preferably integrated into the propellant tank. Therein, the warm propellant can pass on its stored heat thereby preferably evaporating additional propellant. The additionally evaporating propellant may serve for maintaining tank pressure.

After flowing through the second heat exchanger the propellants may enter the combustor, where they produce heat through chemical reaction. This design eliminates moving parts like (turbo-)pumps, promising higher reliability and lower weight. Additionally, no high-pressure gas tank is required to control pressure in the propellant tanks. The phase (gaseous or liquid) of the propellant reaching the thruster can be controlled. Phases are preferably kept separate prior to reaching the thruster. Single phase injectors can be supplied with the correct phase, but also the composition of two-phase mixtures can be precisely controlled. Since each propellant tank is heated by a branch of the propellant it contains, condensation in the heat exchanger is reduced to a minimum or even eliminated.

The measures disclosed herein are suitable and designed for selfpressurizing, possibly cryogenic, propellants. No pumps, high-pressure tanks or pressurization gas are required. Both oxidizer and fuel may be used for thruster (e.g. combustor) cooling. Gaseous propellants can be extracted from tanks (at least when the thruster is active). There may be a starting time frame in which both liquid and gaseous propellants are fed to the thruster inadvertantly, however as soon as the thrust is large enough, the liquid and gaseous phase separate in the propellant tanks and can be extracted separately. Waste-heat from the thruster (e.g. engine) is used to maintain self-pressurization.

A monopropellant system or multi-propellant systems are possible. In the latter case the other branch(es) may be of conventional type - with turbo pump or high-pressure gas tank - or follow a layout similar to the shown branch. The propellant is withdrawn from the tank in gaseous form. Then it is used to cool the thruster in counter-flow, but other flow configurations are possible.

Depending on the size and power setting of the thruster, not all heat may be needed to sustain the tank pressure. Valves in the heat exchanger line and in the bypass line allow to control the tank pressure. Depending on the requirements of the thruster, the valves may be simple magnetic valves that alternatingly open and close their respective branches, or more sophisticated control valves allowing a continuous adjustment of mass flow rate and pressure drop.

The two valves may also be replaced by a single 3-way valve at the bypass outlet. Depending on the requirements of the thruster, a continuously adjustable valve may be necessary or preferable. A 3-way valve may also be placed upstream from the heat exchanger instead. In another configuration there is only one valve in the bypass line. When it is closed all propellant passes the heat exchanger. When it is (fully) open, most of the propellant is routed through the bypass due to its lower pressure loss. If, for open bypass valve, the ratio between the two mass flow rates through heat exchanger and bypass does not meet the requirements, it can be adjusted by placing an orifice in one of the branches, or both.

The valve may be placed in the heat exchanger branch instead (not shown). Such a setup will however most probably require some throttling device, e.g. an orifice, in the bypass line, which is then always open.

A part of the propellant may also be withdrawn from the tank in liquid form. It is then directly routed to the thruster. Depending on the size and power of the thruster, cooling of the walls may not yield enough heat to evaporate all propellant consumed. If in such a situation all propellant was withdrawn from the ullage, i.e. in gaseous form, the tank pressure could not be sustained. Withdrawing propellant in liquid form requires significantly less heat input to sustain the tank pressure. The phases of the propellant remain separated. If a part of the thruster requires gaseous or liquid propellant only it can be fed from the respective line. This setup essentially lifts the thrust limitation of a gaseous injection rocket.

The propellant can also be used for a cold gas system for attitude control. Thereby the otherwise required high-pressure gas tank is eliminated. If the propellant is a monergol, attitude control can be achieved with more efficient hot firing thrusters. The attitude control system can withdraw its propellant from the propellant line, e.g. downstream from the cooling channels. This would allow preheating when the main engine is operated at the same time, which is more efficient.

In case of bipropellant engines, each propellant tank is heated by returning its own heated-up propellant. Therefore, condensation in the heat exchanger is nearly impossible and the delivery of purely gaseous propellant guaranteed. The independence of the different propellant branches from one another eliminates any essential difficulty from extending the Monopropellant system to a bipropellant system and to a multi-propellant system of any number of propellants. When designing a thruster for a bipropellant or multi-propellant system, the cooled surface of the thruster must be divided between the propellant branches, such that each branch picks up enough heat to maintain tank pressure.

Propellant branches may be coupled by a heat exchanger. This is preferably, if one propellant is incompatible with the thruster material or may not be heated up arbitrarily. Depending on the requirements, a heat exchanger can couple the secondary propellant branch to various parts of the primary propellant branch, e.g. downstream from the mixing section, in the primary heat exchanger return line or in the thruster cooling return line.

Depending on the requirements, a heat exchanger may be of co-flow (lightweight, limited heat pick-up), counter-flow (efficient, but heavy) or any other type.

BRIEF DESCRIPTION OF THE DRAWINGS

The invention will be described with reference to the accompanying schematic drawings that are listed below.

Fig. 1 depicts a first embodiment of a propulsion system;

Fig. 2 depicts a second embodiment of a propulsion system;

Fig. 3 depicts a third embodiment of a propulsion system;

Fig. 4 depicts a fourth embodiment of a propulsion system;

Fig. 5 depicts a fifth embodiment of a propulsion system; and

Fig. 6 to

Fig. 10 depict embodiments of a valve arrangement used in any of the propulsion systems.

DETAILED DESCRIPTION OF PREFERRED EMBODIMENTS

Referring to Fig. 1 , the propulsion system comprises a main thruster 01 . The main thruster 01 is a combustor-type thruster.

The system comprises a primary propellant tank 101 (propellant tank) that is filled with a primary propellant. The primary propellant is fed from the propellant tank 101 in gaseous form through a primary gaseous outlet (gaseous outlet arranged facing away from the main thruster 01 such that the ullage is at the location of the outlet under acceleration) and a primary gaseous feedline 102 to a primary branch thruster heat exchanger 105 (first heat exchanger). The first heat exchanger 105 is arranged at the main thruster 01 in counter-flow configuration. Other flow configurations are possible. The primary propellant is heated by the waste heat of the thruster 01 .

After the first heat exchanger 105, the heated propellant passes through a primary valve arrangement 106 (valve arrangement).

The valve arrangement 106 has a plurality of branches and at least one valve device.

The valve arrangement 106 may include a first branch A-B that feeds the output of the first heat exchanger 105 to a primary branch tank heat exchanger 107 (second heat exchanger). The second heat exchanger 107 is arranged within the propellant tank 101 and heats the remaining propellant, so as to maintain a predefined pressure.

The valve arrangement 106 may include a second branch A-D that feeds the output of the first heat exchanger 105 to a gaseous injector feed line 108.

The valve arrangement 106 may include a third branch C-D that fluidly connects the output of the primary branch tank heat exchanger 107 to the gaseous injector feed line 108. The propellant is fed from the second heat exchanger 107 into the gaseous injector feedline 108 via the third branch C-D in a gaseous state.

It should be noted that in normal operation the propellant remains gaseous after being discharged from the gaseous outlet and throughout its entire flow path in the propellant cycle.

The gaseous propellant is then injected through the gaseous injector feed line 108 into the main thruster 01 .

In the following, embodiments are only described insofar as the differ from the previously described embodiment.

As depicted in Fig. 2, in addition a liquid bypass is provided. The liquid bypass includes a primary liquid control valve 104 (liquid control valve) and a primary liquid feedline 103 (liquid feedline). When the main thruster 01 is active, liquid propellant is discharged through a liquid outlet at the side of the propellant tank 101 that is closer to the main thruster 01 and fed through the liquid control valve 104 and the liquid feedline 103 into the main thruster 01.

As shown in Fig. 3, in addition an auxiliary thruster 02 is provided. The auxiliary thruster 02 may be used for attitude control, for example. The auxiliary thruster 02 may be a cold-gas thruster or hot fire. The auxiliary thruster 02 is fed from the primary gaseous feedline 102 (or the first heat exchanger 105, not shown) through a primary attitude control feedline 109. A primary attitude control main valve 110 is provided for controlling the auxiliary thruster 02. It should be noted that, while not explicitly shown here, depending on the overall configuration, the auxiliary thruster 02 may also be a bi- or multipropellant thruster.

As shown in Fig. 4, the system may be configured for bipropellant. The system has a primary flow circuit which includes the elements starting with a 1 and a secondary flow circuit which includes the elements staring with a 2. For sake of brevity, the secondary flow circuit is not described in more detail as it is identical to the primary flow circuit in the first embodiment, to which reference is made. It should be noted that, while not explicitly shown here, the system may be configured for multipropellant operation, i.e. using three or more propellants.

As shown in Fig. 5, a branch heat exchanger 03 is provided. The branch heat exchanger 03 is arranged in the primary flow circuit between output D of the valve arrangement 106 and the gaseous injector feedline 108. The branch heat exchanger 03 is arranged in the secondary flow circuit between the secondary propellant tank 201 on the secondary gaseous feed line 202 and input A of the secondary valve arrangement 206. In a variant (not shown) the branch heat exchanger 03 is preferably arranged in the primary flow circuit between the output of the second heat exchanger 107 and input C of the primary valve arrangement 106. In this variant the branch heat exchanger 03 is arranged in the secondary flow circuit between the secondary propellant tank 201 on the secondary gaseous feed line 202 and input A of the secondary valve arrangement 206.

It should be noted that in this embodiment, preferably the secondary flow circuit has no secondary branch thruster heat exchanger 205.

Generally, in all embodiments, the gaseous propellant is fed from the ullage in the propellant tank 101 through the primary gaseous feed line 102 to the first heat exchanger 105 at the main thruster 01 . From there the gaseous propellant is fed into the valve arrangement 106 which distributes the gaseous propellant between the second heat exchanger 107 (first branch A-B) and the direct feed to the main thruster 01 (second branch A-D).

Referring to Fig. 6 through Fig. 10, different embodiments of the valve assembly 106 are described in more detail.

As shown in Fig. 6, the second branch A-D and the third branch C-D are without a valve. The first branch A-B includes a control valve cu2 that controls how much gaseous propellant is fed to output B, i.e. to the second heat exchanger 107, whereby the flow is directly controlled. The flow through the primary bypass cu1 towards output D and the main thruster 01 is indirectly controlled. In this configuration the control valve cu2 is also called a primary heat exchanger control valve.

As shown in Fig. 7, the first branch A-B and the third branch C-D are without a valve. The second branch A-D includes a control valve cu3 that controls a primary bypass cu1 . By controlling the control valve cu3, the amount of gaseous propellant that is fed to output B, i.e. to the second heat exchanger 107, is indirectly controlled. The flow through the primary bypass cu1 is directly controlled. In this configuration the control valve cu3 is also called a primary bypass control valve.

As shown in Fig. 8, only the third branch C-D is without a valve. Both the first branch A-B and the second branch A-D include a control valve cu2 and cu3 respectively. The flow of gaseous propellant to the second heat exchanger 107 and directly to the thruster 01 can be controlled independently from each other. In this configuration the control valves cu2, cu3 are designated as described before.

As shown in Fig. 9, the first branch A-B is without a valve. The second branch A-D and the third branch C-D include a common control valve cu4. The control valve cu4 thereby controls indirectly the flow through the second heat exchanger 107 and the primary bypass cu1 . In this configuration the common control valve cu4 is also called primary mixing valve.

As shown in Fig. 10, the third branch C-D is without a valve. The first branch A-B and the second branch A-D include a common control valve cu5. The control valve cu5 thereby controls directly the flow through the second heat exchanger 107 and the primary bypass cu1 . In this configuration the control valve cu5 is also called primary splitting valve.

List of reference signs:

01 main thruster

02 auxiliary thruster

03 branch heat exchanger

101 primary propellant tank

102 primary gaseous feedline

103 primary liquid feedline (liquid feedline)

104 primary liquid control valve (liquid control valve)

105 primary branch thruster heat exchanger (first heat exchanger)

106 primary valve arrangement (valve arrangement)

107 primary branch tank heat exchanger (second heat exchanger)

108 gaseous injector feed line

109 primary attitude control feedline

110 primary attitude control main valve I

201 secondary propellant tank

202 secondary gaseous feed line

205 secondary branch thruster heat exchanger (first heat exchanger)

206 secondary valve arrangement

207 secondary branch tank heat exchanger (second heat exchanger)

208 gaseous injector feed line

A-B first branch

A-D second branch

C-D third branch cu1 primary bypass cu2 control valve cu3 control valve cu4 common control valve cu5 common control valve

Claims

1 . A, preferably pumpless, pressure fed propulsion system for a spacecraft, preferably a satellite or microsatellite or space tugs, the system comprising: a propellant tank (101 , 201 ) for storing a propellant; a first heat exchanger (105, 205) that is fluidly connected to the propellant tank (101 , 201 ) and arranged downstream thereof so as to be able to heat the propellant with waste heat from a thruster (01 ); a second heat exchanger (107, 207) that is fluidly connected to the first heat exchanger (105, 205) and arranged downstream thereof so as to be able to heat the stored propellant; and a thruster (01 ) that is fluidly connected to the second heat exchanger (107, 207) and arranged downstream thereof so as to be able to generate thrust using the propellant.

2. The system according to claim 1 , wherein the propellant tank (101 , 201 ) comprises a gaseous outlet that is configured such that when the thruster (01 ) is active only gaseous propellant is discharged from the gaseous outlet.

3. The system according to any of the preceding claims, further comprising a valve assembly (106) that is configured to control a gaseous propellant flow from the first heat exchanger (105, 205) to the second heat exchanger (107, 207) and/or to the thruster (01 ).

4. The system according to claim 3, wherein the valve assembly (106) comprises at least one of the following: a heat exchanger control valve (cu2) that is arranged to control a direct gaseous propellant flow from the first heat exchanger (105, 205) to the second heat exchanger (107, 207); and/or a gaseous bypass valve (cu3) that is arranged to control a direct gaseous propellant flow from the first heat exchanger (105, 205) to the thruster (01 ); and/or a mixing valve (cu4) that is arranged to receive a gaseous propellant flow directly from the first and second heat exchangers (105, 205; 107, 207) and controllably discharge a gaseous propellant flow mixed therefrom to the thruster (01 ); and/or. a splitting valve (cu5) that is arranged to receive a gaseous propellant flow directly from the first heat exchanger (105, 205) and controllably discharge a split gaseous propellant flow directly to the second heat exchanger (107, 207) and to the thruster (01 ).

5. The system according to any of the preceding claims, wherein the propellant tank (101 , 201 ) includes a liquid outlet that is arranged such that when the thruster (01 ) is active only liquid propellant is discharged from the liquid outlet, and the system further comprises a controllable liquid bypass (103) that fluidly connects the liquid outlet to the thruster (01 ) so as to feed liquid propellant to the thruster (01 ) in a controllable manner.

6. The system according to any of the preceding claims, further comprising an auxiliary thruster (02) that is fluidly connected to the propellant tank (101 ), preferably a gaseous flow from the propellant tank, in a controllable manner.

7. The system according to any of the preceding claims, wherein the propellant tank is a primary propellant tank (101 ) for storing a primary propellant; the first heat exchanger is a primary branch thruster heat exchanger (105) that is fluidly connected to the primary propellant tank (101 ) and arranged downstream thereof so as to be able to heat the primary propellant with waste heat from the thruster (01 ); the second heat exchanger is a primary branch tank heat exchanger (107) that is fluidly connected to the primary branch thruster heat exchanger (105) and arranged downstream thereof so as to be able to heat the stored primary propellant; and the system further comprises: a secondary propellant tank (201 ) for storing a secondary propellant; a secondary branch thruster heat exchanger (205) that is fluidly connected to the secondary propellant tank (201 ) and arranged downstream thereof so as to be able to heat the secondary propellant with waste heat from the thruster (01 ); a secondary branch tank heat exchanger (207) that is fluidly connected to the secondary branch thruster heat exchanger (205) and arranged downstream thereof so as to be able to heat the stored secondary propellant; and the thruster (01 ) is fluidly connected to the primary and secondary branch tank heat exchangers (107, 207) and arranged downstream thereof so as to be able to generate thrust using the primary and secondary propellants.

8. The system according to claim 7, wherein the primary propellant tank (101 ) comprises a primary gaseous outlet that is arranged such that when the thruster (01 ) is active only gaseous primary propellant is discharged from the primary gaseous outlet, and the secondary propellant tank (201 ) comprises a secondary gaseous outlet that is arranged such that when the thruster (01 ) is active only gaseous secondary propellant is discharged from the secondary gaseous outlet.

9. The system according to any of the claims 7 or 8, wherein the valve assembly is a primary valve assembly (106) that is configured to control a gaseous primary propellant flow from the primary branch thruster heat exchanger (105) to the primary branch tank heat exchanger (107) and/or to the thruster (01 ), and further comprising a secondary valve assembly (206) that is configured to control a gaseous secondary propellant flow from the secondary branch thruster heat exchanger (205) to the secondary branch tank heat exchanger (207) and/or to the thruster (01 ).

10. The system according to any of the claims 7 to 9, wherein the liquid outlet is a primary liquid outlet that is arranged such that when the thruster (01 ) is active only liquid primary propellant is discharged from the primary liquid outlet, and the liquid bypass (103) is a primary controllable liquid bypass that fluidly connects the primary liquid outlet to the thruster (01 ) so as to feed liquid primary propellant to the thruster (01 ) in a controllable manner; wherein the secondary propellant tank (201 ) includes a secondary liquid outlet that is arranged such that when the thruster (01 ) is active only liquid secondary propellant is discharged from the secondary liquid outlet, and the system further comprises a controllable secondary liquid bypass that fluidly connects the secondary liquid outlet to the thruster (01 ) so as to feed liquid secondary propellant to the thruster (01 ) in a controllable manner.

11 . The system according to any of the claims 7 to 10, further comprising a branch heat exchanger (03) that is arranged to exchange heat between a gaseous primary propellant flow and a gaseous secondary propellant flow.

12. The system according to claim 11 , the branch heat exchanger (03) is arranged in a primary propellant circuit between the primary branch tank heat exchanger (107) and the thruster (01 ) and in a secondary propellant circuit between the secondary propellant tank (201 ) and the secondary branch tank heat exchanger (207), preferably a secondary valve arrangement (206).

13. A spacecraft comprising a propulsion system according to any of the preceding claims.

14. A method for pressure feeding a thruster (01 ) of a propulsion system for a spacecraft, preferably a satellite or microsatellite or a space tug, the method comprising: feeding gaseous propellant from a propellant tank (101 , 201 ) to a first heat exchanger (105, 205); with the first heat exchanger (105, 205) heating the propellant with waste heat from a thruster (01 ) and feeding the heated propellant to a second heat exchanger (107, 207); with the second heat exchanger (107, 207) heating the remaing propellant in the propellant tank (101 , 201 ) and subsequently feeding the propellant from the second heat exchanger (107, 207) to the thruster (01 ), and generating thrust with the propellant.