WO2024017865A1 - Aerospace vehicle having a spike engine, and methods of operating and simulating thereof - Google Patents

Abstract

Aerospike engine portion, comprising: - a tube like housing including an outer wall, - a preferably substantially conical spike part arranged at the lower side of the housing in the exhaust of the engine part - fluid lines extending from the upper side of the engine part to the lower side into the spike part and extending vice versa from the spike part to the upper side of the engine part.

Description

AEROSPACE VEHICLE HAVING A SPIKE ENGINE, AND METHODS OF OPERATING AND SIMULATING THEREOF

Aerospace vehicles with aerospike nozzles were explored during the 1960's. After the project X-33 by Lockheed Martin was abandoned in the early 2000' s, there has been little effort for operational or flight use.

Current rocket engines are typically "bell shaped"; wherein an outside physical bell contains the combustion chamber and directs the thrust of the outgoing gasses. These traditional engines add weight to the design, also because of the necessary steering mechanism for the bell shaped shell. For bigger rockets usually more stages are necessary. For smaller rockets the design is not very fuel efficient. It is difficult to render aerospace more sustainable, because of the size and complexity of this type of engine, also in view of the desired re-use of components.

To overcome the size/weight/payload issue, typical rocket systems use multiple stages, where the first stage engine is decoupled from the rocket with payload after the first stage, allowing the second stage to continue with less mass. These extra stages increase mission complexity and failure risk.

To direct the output flow of the bell shaped engine (and thereby steering the entire rocket system), parts such as gimbals have to be moved with great force to direct the bell to steer the rocket.

The following documents were retrieved during a patent search: US 10527003 B, EP 3597897 A , WO 2018/045351 A, WO 2015/155733 A 5, and US 5221045 A, which documents appear to be insufficiently relevant for the present design.

The present patent disclosure primarily provides an aerospike engine portion, comprising: a tube like housing including an outer wall, a substantially conical spike part arranged at the lower side of the housing in the exhaust of the engine part a number of fluid lines extending from the upper side of the engine part to the lower side into the spike part and extending vice versa from the spike part to the upper side of the engine part.

Preferably propellant (oxidizer and/or fuel) flows through the fluid lines and then returns to the upper part where is decomposes, preferably using honeycomb inserts.

By preferably integrating some fluid lines in the walls of the spike part some weight loss is also achieved. Further fluid lines are preferably centrally located in the tube like housing.

Preferably the aerospace vehicle is provided with a first tank in which hydrogen peroxide is stored and a second tank wherein RP-1 (Rocket Propellant) is stored; both liquid propellants are also configured for cooling the spike part when pumped through the lines. Preferably an electric pump is used.

A simulation has shown that overheating is prevented by using the above fuels and the cooling lines in the interior of the spike part.

In a first preferred design an aerospace vehicle can bring a payload of about 5 kg at an height of about 50 km, the rocket engine showing a thrust of about 5,000 N.

In another preferred design the rocket having a thrust e.g. of about 50,000 N am much heavier payload can be brought in outer space, while the engine is also configured for reentry in the atmosphere.

The present disclosure will be illustrated along the following description of a detailed disclosure in which details, features and advantages of the design and in which reference is made to a drawing, in which show: fig. 1 a side view, partly schematic, and partly broken away, of a section of the design; and fig. 2 a section over line a-b in fig. 2.

Cooling of the aerospike engine is one of the main objects of the present design. Cooling of a spike nozzle far more challenging than cooling of a bell-shaped nozzle. Increasing the thickness of the combustion chamber walls, imposing a large mass penalty, is going into the wrong direction.

The aerospike nozzle's performance advantages come from the complex geometry it encompasses. The optimal method for producing such complex geometry can be effectively achieved through additive manufacturing (metal 3D printing).

The present design provides the following advantages: increased efficiency of 10-15 % over all altitudes; smaller size; lower drag; un-gimballed steering; no stages; lower failure risks.

In figure 1 an aerospike engine 5 is mounted in an rocket assembly 1 in an housing 2 of which a rocket payload is mounted as well as some rocket and engine control systems. Oxidizer tank 3 and fuel tank 4 use a pressure medium , preferably Helium, at 0.3MPa to flow the oxidizer and fuel into the pump mechanism of the tank.

High Test Peroxide (HTP) is preferably used as oxidizer. An oxidizer pump flows HTP out at a substantially constant flow rate 1.586kg/s at 4-5MPa into an oxidizer channel 6.

The oxidizer enters from the top of engine through channel 6 in downward direction (when the assembly is standing or flying substantially vertically in upward direction) into centrally located spike cooling channels at room temperature, proceeding through channels 7 and flowing upward again to reach cooling manifold 8. The oxidizer exits spike cooling channels at a temperature of 370K, or lower, into oxidizer manifold 8 at the top of the engine. The oxidizer is distributed by the oxidizer manifold and injected at 3.3MPa pressure into decomposition chamber 9.

The oxidizer undergoes a decomposition process using a catalyst, preferably including circular honeycomb pads, and is converted to superheated steam and gaseous oxygen at about 1000K in chamber 10, under a pressure of 3.2MPa. In chamber 10 the oxidizer will be combined with fuel injected in the combustion chamber 15, at a pressure of 2.6MPa.

Gaseous byproducts flow through tubes that pass through the fuel manifold (without further mixing) and are injected at 2.6MPa pressure into combustion chamber 15.

RP-1 is preferably used as fuel. A fuel pump moves the RP-1 out at a substantially constant flow 0.214kg/s at 4-5MPa into the fuel channel 11.

The fuel enters the cooling channels (at 12) at room temperature. The cooling channels 13 are integrated in the outside wall of the combustion chamber. The fuel moves down and goes back up again through a number channels.

In figure 2 a cross-section of the channels at Line AB is shown.

The fuel exits cooling channels 13 at a temperature of 450K or less into fuel manifold 14. The fuel is distributed throughout the fuel manifold and injected at 2.6MPa pressure into combustion chamber 15.

The combined fuel and oxidizer in the combustion chamber result in a thrust 16.

In a perspective view 9 (fig.2) of a cross-section of the cooling channels in the combustion chamber the outer surface 1 of the chamber is shown; the inner area 3 is the combustion chamber. The cross section shows the relative small size of cooling channels 4. Each channel can contain cooling liquid (fuel) moving up or down (perpendicular to the surface of the figure). Both ends of the channels can be connected with U-shaped connections, as designed. In this way an efficient cooling flow pattern can be realized using the entire chamber without direct contact of fuel and hot combustion gasses.

The design has been subject to simulations which have shown that during launch and flight of a small space vehicle the temperatures stay below critical limits for the functioning of the structural and other operating parts of the space vehicle. Actual tests will also have to be successful before the aerospike design will finally take over the design of rockets, such that efficiency is substantially increase as the thrust is adjusted by the ambient pressure. Also big rockets can then be designed with a single stage.

The requested rights are defined by the annexed claims within the scope of which many variations are conceivable by one skilled in the art, also beyond the above description of a preferred design.

Claims

1 Aerospike engine portion, comprising: a tube like housing including an outer wall, a preferably substantially conical spike part arranged at the lower side of the housing in the exhaust of the engine part fluid lines extending from the upper side of the engine part to the lower side into the spike part and extending vice versa from the spike part to the upper side of the engine part.

2. Aerospike engine part according to claim 1, wherein the fluid lines extending upwardly debouche into the interior of upper part of the tube like outer wall of the housing.

3. Aerospike engine part of claim 1 or 2, wherein the interior of the upper part of the outer wall is provided with honeycomb shaped inserts which are configured to decompose propellant ( fuel and/or oxidizer)

4. Aerospike engine part of claim 1,2 or 3, wherein the fluid lines comprise channels in the wall of the conically shaped spike part.

5. Aerospike engine part of any of claims 1-4, wherein the fluid lines comprise conduits which are centrally located in the tube like housing.

6. Aerospace vehicle, comprising an aerospike engine part according to any of claims 1-5, and a first tank configured to hold oxidizer, preferably hydrogen peroxide and connected to a number of the fluid lines.

7. Aerospace vehicle of claim 6 provided with a second tank configured to hold kerosene, preferably RP 1, and connected to other fluid lines.

8. Aerospace vehicle according to any of claims 5,6 or 7, provided with an electric pump configured to pump the oxidizer and/or fuel through the lines.

9. Method of operating an aerospace vehicle according to any of claims 6-9, wherein the engine is cooled by flowing propellant though fuel lines during a predetermined time after ignition .

10. Method of operating an aerospace vehicle according to claim 9, wherein the predetermined time is about two minutes for en engine thrust level of 5,000 N having payload of about 5 kg bringing the aerospace vehicle of about 0.5 meter in height and a diameter of about 30 cm to an attitude of 50 km.

11. Method of operating an aerospace vehicle according to claim 9, wherein the engine thrust is about 50,000 N and the vehicle is configured for travel in outer space and reentry into the earth's atmosphere.

12. Method of simulating the operation of the aerospace vehicle of claim 10, wherein the temperature of all the parts esp. the aerospike engine portion is calculated and visualized by calculations on a computer.