|Title of Invention||
AN EXPANDER CYCLE ROCKET ENGINE
|Abstract||AN EXPANDER CYCLE ROCKET ENGINE|
|Full Text||FORM 2
THE PATENTS ACT 1970
[39 OF 1970]
[See Section 10; Rule 13]
AN EXPANDER CYCLE ROCKET ENGINE
VOLVO AERO CORPORATION, a Swedish Body Corporate, of S-461 81 Trollhattan, Sweden,
The following specification particularly describes the nature of the invention and the manner in which it is to be performed:-
APPARATUS FDR CONTROLLING THE HEAT TRANSFER TO THE NOZZLE WALL OF EXPANDER CYCLE ROCKET ENGINES
The invention relates to an apparatus for increasing the power
of expander cycle rocket engines, particularly to an apparatus
for increasing the heat transfer to the coolant on the inside
of a nozzle wall provided with coolant channels of expander
cycle rocket engines.
By, for instance, US-3 712 546 it is known to control the boundary layer at the nozzle wall ox the rocket engine. The object of this controlling of the/boundary layer is to reduce the friction between the combustion gases and the nozzle wall so it is possible to increase the expansion ratio of rocket nozzles, particularly vacuum nozzles, and thereby the power of the rocket engine. The poorer increase of the rocket engine ac-cording to this document is thus achieved in another way than the apparatus according to the invention.
US-5 363 645 describes an apparatus for transpiration cooling of the combustion gases in the throat of a rocket engine com¬bustion chamber. However, the object Of this invention is to optimise the consutoption of coolant and not to increase the power of the rocket engine.
Usually, rocket engines are divided into different engine cycles depending on how the flow of the oxidant and the fuel are organised in the engines. In rocket engines of stage com-bustion cycle and gas generation cycle the combustion takes place in two places of the engine, i.e. in the main combustion chamber and in a. secondary burner. The contbustion gases from the secondary burner is "used to drive the turbines of the fuel and oxidizer pumps.
However, in expander cycle rocket engines the combustion takes
place only in the main combustion chamber, and the turbines
for driving the fuel and oxidizer pumps are driven by the fuel
(usually liquid hydrogen) flowing through the coolant channels
in the. walls of the main combustion chamber and the nozzle.
Thus, this means that the fuel is fed from the tanks, through
the pumps where the fuel pressure is increased, and through
the coolant channels in the walls of nozzle and the combustion
chamber and then to the turbines of the fuel and oxidizer
pumps and then out into the combustion chamber in which it is
burnt together with oxidizer. This means that the more the
fuel is heated snd expanded the more power can be gained from
the fuel for driving the turbines, whereby the efficiency of
the engine being increased.
The maximum reachable combustion chamber pressure is thus set by how touch the fuel is heated in the coolant channels. There for, it is desirable to obtain as high combustion chamber pressure as possible, since this will give the largest power of the rocket engine.
To increase the pressure and thereby the power of an expander cycle rocket engine it is extremely important to maximize the
heat transfer to the fuel for increasing its temperature. Even
a small increase of the temperature of "the fuel has great im-
portance, since the power of the engine will thereby be in-
heat transfer to the fuel comprises i.a. increasing of the area of the nozzle wall facing the flame, for instance, by forming the nozzle wall of pipes with half circular or circu¬lar cross-section. Another way is to maks the nozzle wall of a material with^high heat conductivity such as copper.
The drawback of the known technique in the area of the inven-tion is that nozzle walls formed of pipes with half-circular or circular cross-section has low strength in the djrecticn of the tangent and must therefore be reinforced on the outside with different means. This means that the nozzle will be heavy and payload capacity is lost. To make a nozzle wall of copper has the drawback, that copper is difficult to weld and has lower tensile strength than, for instance, steel and nickel based materials which means that a nozzle of copper will be heavier than a corresponding nozzle of steel.
The object of the invention is to eliminate the above-mention ed drawbacks of the prior art.
This object is achieved according to the invention in that, to disturb the boundary layer at the nozzle wall and thereby in¬crease the heat transfer, the inside of the nozzle wall facing the flame has a particularly chosen increased surface rough-ness of such a magnitude that it penetrates the viscous sub¬layer of the boundary layer at the nozzle wall.
A non-limiting example of the invention will now be described with reference to the acconipanying drawing which is a. cross-section view of one half of a rocket nozzle with attached com-bustion chamber according to the invention.
As can be seen from the drawing the inside of the nozzle wall 1 shows a particularly chosen increased surface x-oughness. This surface roughness must be so large that it penetrates the viscous sub-layer of the boundary layer.
A lower limit in which the surface roughness will have full effect on the heat transfer can be defined as follows:
where cf is skin friction, u is the velocity at tne boundary layer, v is the viscosity and y is the distance normal to the
The surface roughness should be at least 50 y+ which is the lower limit within which the surface roughness without any doubt will penetrate the viscous sub-layer of . the boundary layer.
For a typical Nozzle of art expander cycle rocket engine the surface roughness should increase progressively front the inlet to the outlet of the nozzle. At the inlet the surface rough¬ness should be about 0.15 mm and at the outlet about 1 ram.
This surface roughness on the inside of tne nozzle^ can be achieved by, for instance, machining, such as grinding, mill-ing or by depositing of material through flame or plasma spraying.
By increasing the surface roughness in this way it is possible to achieve an increase of the tentpcrature of the coolant (fuel) of at least 10 K which produces an increase of the cooling effect of at least 1% or more of the rocket engine of the above-mentioned type.
1. An expander cycle rocket engine comprising a rocket nozzle with combustion chamber characterized in that the inside of said nozzle is provided with coolant channels to disturb the boundary layer at the nozzle wall and thereby increasing the heat transfer, the inside of the nozzle wall facing the flame is provided with a particularly chosen increased surface roughness of such a magnitude that it penetrates the viscous sub-layer of the boundary layer at the nozzle wall.
2. An expander cycle rocket engine as claimed in claim 1, wherein the surface roughness increases progressively from the inlet to the outlet of the nozzle.
3. An expander cycle rocket engine as claimed in claim 2, wherein the surface roughness varies from 0,15 mm at the inlet of the nozzle to 1 mm at the outlet of the nozzle.
Dated this 11th day of October, 2001.
of Remfry & Sagar
Attorney for the Applicants
|Indian Patent Application Number||IN/PCT/2001/01255/MUM|
|PG Journal Number||20/2007|
|Date of Filing||11-Oct-2001|
|Name of Patentee||VOLVO AERO CORPORATION|
|Applicant Address||A SWEDISH BODY CORPORATE , OF S-461 81 TROLLHATTAN, SWEDEN.|
|PCT International Classification Number||N/A|
|PCT International Application Number||N/A|
|PCT International Filing date||2001-03-16|