Lutz
KAYSER
OTRAG Rocket
OTRAG Rocket
http://en.wikipedia.org/wiki/OTRAG
OTRAG
OTRAG (German: Orbital Transport und Raketen AG, or Orbital Transport and Rockets, Inc.), was a German company based in Stuttgart, which planned in the late 1970s and early 1980s to develop an alternative propulsion system for rockets. OTRAG was the first commercial developer and producer of space launch vehicles. The OTRAG Rocket claimed to present an inexpensive alternative to existing launch systems through mass-production of Common Rocket Propulsion Units (CRPU).
History
OTRAG was founded in 1975 by the German aerospace engineer Lutz Kayser. Its goal was to develop, produce, and operate a radically different, low cost, satellite launch vehicle.
The OTRAG rocket was intended to be an inexpensive alternative to the European rocket Ariane and the NASA space shuttle.[1] Kayser and a private consortium of six hundred European investors financed the development and production of the OTRAG satellite launch vehicle. Dr.Ing Kurt H. Debus served as Chairman of the Board of OTRAG after his retirement as director of NASA's Kennedy Space Center,[2] and Dr. Wernher von Braun served as scientific adviser to Kayser.
In the face of doubts by Debus and von Braun, Kayser chose in 1975 to set up testing and launch facilities in Shaba, Zaire (now Katanga Province in the Democratic Republic of the Congo). Debus and von Braun were concerned about the possibility of Zairian acquisition of missile technology from the facilities. Kayser decided to proceed despite their opposition, and testing began at the site in 1977.
Political pressure to halt the company's operations mounted quickly. France and the Soviet Union were historically opposed to German long-distance rocket development, and pressured the Congolese government into closing down the development facility in 1979. Immediately afterwards, Presidents Giscard d'Estaing of France and Leonid Brezhnev of the Soviet Union convinced the West German government to cancel the OTRAG project and close down its German operations. In 1980, OTRAG moved its production and testing facilities to a desert site in Libya. A series of successful tests were conducted at this site beginning in 1981.[3]
Rocket design
OTRAG was a design quite different from conventional multi-stage rockets. The OTRAG design used parallel stages assembled from parallel tanktubes with flat bulkheads. The rockets were designed to carry loads up to two tons, the then usual weight of a communications satellite, into a geostationary orbit. It was planned to later increase the capacity to ten tons or more using multiple identical modules.
The rocket was to consist of individual pipes, each 27 cm in diameter and six meters long. Four of these pipes would be installed one above the other resulting in a 24 meter long fuel and oxidizer tank with a rocket engine at the lower end. The fuel was intended to be kerosene with a 50/50 mixture of nitric acid and dinitrogen tetroxide as an oxidiser. Ignition was provided by a small quantity of furfuryl alcohol injected before the fuel, which ignites hypergolically (immediately and energetically) upon contact with the nitric acid. To simplify the design, pumps were not used to move the fuel to the engines, instead the fuel tanks were only 66% filled, with compressed air in the remaining space to press propellants into the ablatively cooled combustion chamber. Thrust control is by partially closing the electromechanical propellant valves. Pitch and yaw control can thus be achieved by differential throttling. In principle this is extremely reliable and cheap in mass production.
The modular design was intended to result in a large cost reduction due to economies of scale. The CRPU-based satellite launching rocket was estimated to cost approximately one tenth of conventional designs. Automated production processes for all components would reduce labor cost from 80% to 20% and remove the justification for reusability of spent stages.
Controversies and future outlook
Only a few political controversies are known concerning OTRAG because of concerns of neighbors of Zaire and Libya about the dual use potential of rockets. A full orbital launch vehicle was never assembled. Modules were flight tested in Zaire and Libya. 6,000 static rocket engine tests and 16 single stage qualification tests were made to prove the concept as feasible.
The German minister of foreign affairs at that time, Hans Dietrich Genscher, is said to have finally stopped the project under pressure from France and the Soviet Union, and West Germany joined the co-financed "European rocket" Ariane project, which made the OTRAG project unnecessary and eliminated political entanglements of a still divided Germany in the early 1980s.
NASA Commercial Orbital Transportation Services announcement stipulates a 50% US National ownership in vendor companies like von Braun Debus Kayser Rocket Science LLC, DE (BDKRS). This would force Kayser, a German citizen, to sell at least 50% of his shares to Americans.
More recently, the company has been advising Interorbital Systems, resulting in a similar modular rocket design for their Neptune series.[4]
John Carmack, founder and lead engineer of Armadillo Aerospace has stated in his monthly reports and in forum posts that he expects his path to an orbital vehicle to include modular rockets similar to OTRAG technology. Kayser, being the founding engineer of OTRAG, visited Armadillo in May, 2006 and loaned Carmack some of their original research hardware.
"I have been corresponding with Lutz [Kayser] for a few months now, and I have learned quite a few things. I seriously considered an OTRAG style massive-cluster-of-cheap-modules orbital design back when we had 98% peroxide (assumed to be a biprop with kerosene), and I have always considered it one of the viable routes to significant reduction in orbital launch costs. After really going over the trades and details with Lutz, I am quite convinced that this is the lowest development cost route to significant orbital capability. Eventually, reusable stages will take over, but I actually think that we can make it all the way to orbit on our current budget by following this path. The individual modules are less complicated than our current vehicles, and I am becoming more and more fond of high production methods over hand crafter prototypes." -- June 2006 Armadillo Aerospace Update[5]
External links
http://www.b14643.de/Spacerockets_1/West_Europe/OTRAG/Description/Frame.htm
http://www.bernd-leitenberger.de/otrag.shtml very explanative article about the OTRAG history (de)
http://www.astronautix.com/lvs/otrag.htm (en)
http://www.astronautix.com/astros/kayser.htm
Encyclopedia Astronautica
Lutz
Kayser
Kayser, Lutz T (1939-) German engineer and low-cost rocket pioneer. 1975-1987 developed Otrag concept - clustered large numbers of low-cost storable liquid rocket modules to reduce costs by 10x. Tested in Congo and Libya, but project killed by vested interests.
Kayser started his first serious work as a 17-year old student, founding the Arbeitsgemeinschaft fuer Raketentechnik und Raumfahrt an der Universitaet Stuttgart (Working Group for Rocket Technology and Space Travel at the University of Stuttgart) in 1955. This was the first student group at a German University to be active in rocket propulsion and space flight research - at a time before the first satellites were launched and the feasibility of stable satellite orbits was still disputed in academic circles. Eugen Saenger was the mentor of the group and arranged research grants from the Baden-Wuertemberg Economic Ministry. Kayser received the degree of Diplom-Ingenieur (equivalent to a Master of Science) in Aeronautics and Astronautics from the University of Stuttgart.
Working at the request of the Baden-Wuerttemberg government, Kayser selected the site of the Lampoldshausen Rocket Test Centre. Together with Wolfgang Pilz, he laid out the design for the facility. In the 21st Century it remained the largest such institute in Europe.
Kayser's first major development was a bipropellant satellite attitude control systems (later sold to North American Rocketdyne and the US Air Force Rocket Propulsion Laboratory, Edwards AF Base). He also collaborated with the Rocket Engine Division of the NASA Marshall Space Flight Center in connection with the Saturn-IB clustered engine concept. Kayser developed the first ablative combustion chamber for the H-1 engine, later tested in firings at Huntsville. This was a first step towards his concept of parallel clustering of low-cost ablative engines. Further work on the idea was supported by von Braun and jointly financed by NASA and the German Ministry of Scientific Research. Kayser founded Technologieforschung GmbH (TF) as a commercial spin-off to handle these and other contracts.
Kayser invented, developed, and tested the TIROC rocket engine (Tangential Injection and Rotational Combustion). It was the world's smallest thruster burning Monomethylhydrazine and Nitrogen tetroxide. It delivered 1 newton thrust (0.2 lbf) with a minimum burning time of 1 milliseconds and a demonstrated maximum burning time of 1 million seconds (11 days). The valves had response times of under 1 millisecond and were capable of more than 1 million cycles at a 6 sigma confidence level. Kayser also developed one of the first capillary action gas-liquid separation systems. This guaranteed positive liquid flow from propellant tanks to the rocket engines in zero-gravity. Future applications were high performance satellite and space vehicle attitude control systems.
The first pan-European launch vehicle program was the European Launcher Development Organization's Europa I of the 1960s. This medium lift vehicle consisted of a British first stage (Blue Streak), a French second stage (Veronique), and a German third stage. After several attempts without a single successful launch, the German Government asked Kayser to investigate. Explosions always occurred shortly after cutoff of the French second stage and before ignition of the German third stage. Kayser had just finished assisting Professor Argyris of Stuttgart University in developing the world's first finite element computation method for solution of statics and dynamics of structures (ASKA).
Applying this dynamic simulation to the combined second-third stage ELDO- launch vehicle before separation, and following exhaustive evaluation of all telemetry data Kayser concluded:
...that the rough chamber pressure and thrust oscillations of the second stage rocket engine during cut-off destroyed the intertank bulkhead of the third stage. This in turn hypergolically ignited the liquid propellants (N2O4/N2H4) and the third stage exploded before its engine could even be started.
...the structural design of the stage was fundamentally flawed and would be very expensive to modify.
As a consequence of this report and other performance and political factors, the ELDO program was cancelled in 1972.
Thereafter Kayser's TF received German Government contracts to study and analyse NASA's proposed Space Shuttle project. It soon became clear to Kayser that during the first two years of development too many conflicting US Government requirements were incorporated into the design. He also found that industry desires to sell their respective technologies forced incompatible features together. Solid propellant boosters increased cost. Wings twice as large as necessary for NASA made the delta-winged shuttle orbiter less safe than a lifting body design. Based on Kayser's recommendation the German Government stopped participation in that program. It took NASA 25 years to reach the same conclusion - that the Space Shuttle was inherently flawed and needed a successor as soon as possible.
It was clear that governments had a hard time finding researchers and engineers for impartial assessment of large, expensive and long term space projects. Analysis of such large systems required a very wide knowledge of all scientific fields with decades of experience and an independent view. Kayser had these abilities and as a result became a consultant to NASA, DARPA, USAF, and NRO in formulating future US space programs.
In the early 1970's Willy Brandt's Ministry of Science and Technology solicited a contract for demonstration of launch vehicle technology an order of magnitude cheaper and more reliable than existing launchers. Kayser's research company TF won the contract and developed a radically new rocket technology, making more than 20 inventions in the process.
Kayser's concept involved the parallel clustering of large numbers of identical propellant tank and rocket engine modules. This allowed the application of mass production techniques as used in the automobile industry. This in turn resulted in cost reduction by a factor of 10. This breakthrough and the static testing in of prototype modules at Lampoldshausen stirred concern in the competitive aerospace industry. The established space launch companies were accustomed to making easy guaranteed profits through high cost plus fixed fee government contracts.
In order to exploit this low-cost rocket technology on a commercial basis Kayser founded OTRAG (Orbital Transport und Raketen AG). It was the world's first commercial launcher development, production and launch company.
Wernher von Braun and Kurt Debus, the leading managers of American rocketry, were so enthusiastic about the project that they joined the team after their retirement from NASA. Their contribution was important and helped to introduce lessons learned from earlier programs. Von Braun introduced the concept of parallel clustering of tanks and engines with his Saturn I design and had shown the way towards the low-cost breakthrough 20 years earlier. However, both rocket pioneers were in doubt whether this technology should be flight tested in developing countries because of the possibility that it would be misused for weapons. Kayser optimistically hoped he would be able to limit the technology to commercial satellite launchings. Kayser was proven wrong and suffered heavy losses as a result.
International controversy erupted when Kayser conducted suborbital test flights from launch ranges in the Congo and Libya. 14 suborbital test flights proved the concept and led to a 100% qualification of the technology and the verified the extremely low production cost. However Soviet president Brezhnev and French president Giscard d'Estaing applied heavy political pressure on the German government to stop the project. After a total investment of $ 150 million, OTRAG had to terminate production in Germany. Production was relocated to the launch site in Libya. This in turn led to Libyan military circles eyeing the facilities as a means of obtaining military rocket technology. OTRAG's production and launch range equipment were illegally confiscated, as had happened to the foreign oil industry a decade earlier. All attempts by Kayser to solve the problem were futile. Without Kayser's know-how the Libyans were able to conduct only a few test launches with the stolen equipment. After ten years of desultory testing the Libyan program came to an end.
As of 2005, Kayser was actively searching for partners to fund an OTRAG production facility in the United States and to apply his unique low-cost technology to the requirements of the future American space program. He founded von Braun Debus Kayser Rocket Science LLC to transfer OTRAG's intellectual property and know-how to the United States. Kayser, along with newer private entrepreneurs such as Musk, Rutan, and Bezos, still dream of achieving the goal of affordable space transport below $ 1,000 per pound into orbit.
Birth Place: Stuttgart.
Born: 1939.03.31.
Liquid-fueled
rocket
US3945203
US3945203
A liquid-fueled rocket has a storage tank for the fuel components of a liquid fuel and the storage tank is able to communicate with the combustion chambers of the rocket via normally closed valves. The storage tank is enclosed and the fuel components only in part fill the latter. A body of gas is confined in the storage tank and has a pressure exceeding that in the combustion chambers. Thus, when the valves are opened, the fuel components are caused to flow from the storage tank to the combustion chambers by virtue of the difference between the pressure of the gas and the pressure in the combustion chambers.
BACKGROUND OF THE INVENTION
The invention relates generally to liquid-fueled rockets. More particularly, the invention relates to a method and an arrangement for conveying the liquid fuel of a liquid-fueled rocket from the storage tanks of the rocket to the combustion chambers of the latter.
The fuel components of a liquid-fueled rocket, that is, the combustible component and the oxidizer constituting the liquid fuel, must be forced into the power plant of the rocket by suitable conveying means. In the known rockets of this type, this is accomplished by pumping or by pressurizing the storage tanks for the fuel components.
Various methods have been proposed and used for pressurizing the fuel components in the storage tanks. In these methods, the fuel components are subjected to the action of a pressure gas which latter is, for example, produced by a separate gas generator or a separate solid-gas generator.
The known methods have several disadvantages. The use of pumps is expensive and structurally prohibitive and, in addition, decreases the useful pay load of the rocket because of the weight of the pumps.
Where a pressure gas is used for conveying the fuel components from the storage tanks to the combustion chambers, it is necessary to provide separate containers for the fuel required to generate the pressure gas and it is also necessary to provide a gas generator. In addition, the methods using a pressure gas require an active regulating circuit for regulating the quantity of pressure gas to be introduced into the storage tanks.
In all of the known methods for conveying the fuel components via pressurization, the requisite energy for conveying the fuel components is supplied to the storage tanks from an external source. These methods are expensive and require additional fuel for generating the pressure gas as well as valves, conduits, generators and additional fuel containers. Aside from the structural and operational difficulties associated with the use of these methods, their use causes an increase in the weight of the rocket and, hence, a reduction in the useful pay load.
SUMMARY OF THE INVENTION
It is, accordingly, a general object of the invention to provide a novel method and arrangement for conveying the liquid fuel of a liquid-fueled rocket from the storage tanks of the rocket to the power plant or combustion chambers of the latter.
Another object of the invention is to provide a method and arrangement whereby the liquid fuel of a liquid-fueled rocket may be conveyed from the storage tanks of the rocket to the power plant of the latter in simple manner.
A further object of the invention is to provide a method and arrangement for conveying the liquid fuel of a liquid-fueled rocket from the storage tanks of the rocket to the power plant of the latter without significantly reducing the useful pay load of the rocket.
An additional object of the invention is to provide a method for conveying the liquid fuel of a liquid-fueled rocket from the storage tanks of the rocket to the power plant of the latter which does not require complicated and expensive equipment.
A concomitant object of the invention is to provide an arrangement for conveying the liquid fuel of a liquid-fueled rocket from the storage tanks of the rocket to the power plant of the latter which is simple in its construction and inexpensive.
In accordance with the objects outlined above and others which will become apparent hereafter, the invention provides a method of conveying the fuel components of a liquid-fueled rocket from the storage tanks to the combustion chambers of the rocket wherein fuel components to be conveyed to a combustion space are confined in an enclosed space so as to only partially fill the latter. A quantity of gas sufficient to cause the pressure of the same to exceed the pressure in the combustion space is also confined in the enclosed space. A flow path is established between the enclosed space and the combustion space and this causes the fuel components to flow from the enclosed space to the combustion space by virtue of the difference between the pressure of the gas and the pressure in the combustion space. Also disclosed is a novel arrangement for conveying the fuel components of a liquid-fueled rocket from the storage tanks to the combustion chambers of the rocket.
Thus, according to the invention, the fuel components are forced into the combustion chamber or chambers of the rocket by a gas which is confined under pressure in the same storage tank or tanks as the fuel components themselves. Suitable gases for this purpose are, for example, nitrogen, helium, air or the like.
The initial pressure of the gas, that is, the pressure of the confined gas before the fuel components have begun to flow into the combustion chambers, is dependent upon the type of power plant used for the rocket and may lie, for example, between about 5 and 100 bars. The initial pressure of the gas is favorably between about 20 and 40 bars and, advantageously, is equal to approximately 30 bars. However, the invention is not limited to the values given above. As an example, the initial pressure of the gas may be higher for a rocket which operates in the atmosphere than for a rocket which operates in vacuum.
In a suitable arrangement for carrying out the method according to the invention, at least one of the storage tanks of the rocket includes at least one portion for accommodating the gas and which is in communication with, or may be brought into communication with, the portions of the storage tank which accommodate the fuel components.
Where the storage tanks of the rocket include more than one portion for accommodating the gas, it is possible to interpose suitable valves between these portions and the initial pressure of the gas may be different in the various portions.
Advantageously, the initial volume of the gas, that is, the volume of the confined gas before the fuel components have begun to flow into the combustion chambers, amounts to approximately 20 to 50% of the total volume of the storage tanks.
The portion or portions of the storage tanks accommodating the combustible fuel component may be sealed with a membrane which tears open when the rocket begins to operate or, in other words, when the valves interposed between the storage tanks and the combustion chambers are opened.
The invention not only makes it possible to eliminate special pumps for conveying the fuel components but also makes it possible to eliminate the additional storage tanks, as well as the corresponding conduits and generators, which it would be necessary to provide for the generation of a pressure gas when using those methods wherein the fuel components are conveyed by such a gas and where additional fuel for generation of the latter is required. Furthermore, by utilizing the invention, there is no need to provide special regulating circuits for regulating the quantity of pressure gas introduced into the storage tanks so as to produce the requisite pressure in the latter.
The novel features which are considered as characteristic for the invention are set forth in particular in the appended claims. The invention itself, however, both as to its construction and its method of operation, together with additional objects and advantages thereof, will be best understood from the following description of specific embodiments when read in connection with the accompanying drawings.
BRIEF DESCRIPTION OF THE DRAWING
FIG. 1 is a section through a fuel storage tank of a rocket schematically illustrating an arrangement according to the invention; and
FIG. 2 is a view similar to FIG. 1 but illustrating another arrangement according to the invention.
DESCRIPTION OF THE PREFERRED EMBODIMENTS
Referring to FIG. 1, a storage tank or container for the liquid fuel of a liquid-fueled rocket is indicated generally at 10 and is seen to be of so-called parallel construction. The storage tank 10 includes two portions 12 and 14 for accommodating the combustible fuel component of the liquid fuel and the portions 12 and 14 extend parallel to the longitudinal axis of the storage tank 10. In the illustrated embodiment, the portions 12 and 14 are located diametrically opposite one another. However, instead of being provided with two portions for accommodating the combustible fuel component, the storage tank 10 may be provided with a plurality of such portions, for example, four portions, or may be provided with only one portion for the combustible fuel component. The portions 12 and 14 are only partially filled with the combustible fuel component, for instance, to the level indicated at 15.
The storage tank 10 includes another portion 16, 18, the part 16 of which is filled with the oxidizing fuel component or the oxidizer to the level indicated at 17, that is, the oxidizer only partially fills the portion 16, 18. The part 18 of the portion 16,18 is filled with a body of gas. It will be seen that the storage tank 10 defines an enclosed space in which the fuel components and the gas are confined.
In the illustrated embodiment, the portions 12 and 14 are closed at their upper ends with membranes 19 and 21, respectively, so that the gas in the portion 16,18 does not communicate with the combustible fuel component. However, it is also possible to eliminate these membranes so that the gas in the portion 16,18 is in direct communication with the combustible fuel component. The only purpose of the membranes 19 and 21 is to keep the combustible fuel component separated from the oxidizer during storage and when the rocket is transported. The membranes 19 and 21 may be of a thin synthetic resin foil or a synthetic resin sheet such as, for example, polyester. Below the lower end of the storage tank 10 there are located the propulsion means for the rocket including the combustion space or combustion chambers 20 and 22. Suitable valves 24 are interposed between the combustion chambers 20 and 22 and the portions 12 and 14, as well as the portion 16,18, of the storage tank 10. The valves 24 are normally closed, that is, the valves 24 will usually be closed until the rocket is to be operated, and when the valves 24 are opened a flow path is established between the combustion chambers 20 and 22 and the portions 12 and 14, as well as the portion 16,18, of the storage tank 10.
It is pointed out here that the construction and arrangement of the combustion chambers and the valves, as well as of the conduit system for introducing the fuel components into the combustion chambers, do not form part of the invention per se and, hence, are only schematically illustrated.
In operation of the embodiment of FIG. 1, the portions 12 and 14 of the storage tank 10 are filled with a combustible fuel component to the level 15 and the portion 16,18 is filled to the level 17 with an oxidizer. The part 18 of the portion 16,18 is then filled with a gas such as, for example, nitrogen, in a quantity sufficient that the gas is at a superpressure. The important consideration is that the pressure of the confined gas exceeds the pressure in the combustion chambers 20 and 22. The gas is thus in direct communication with the oxidizer and, if the membranes 19 and 21 are not used, will also be in direct communication with the combustible fuel component.
When the rocket is to be started or fired, the valves 24 are opened as a result of which the pressure in the portions 12 and 14, which are closed by the membranes 19 and 21 in the embodiment being discussed, decreases. This decrease in pressure, in conjunction with the fact that the pressure of the gas in the portion 16,18 exceeds that in the combustion chambers 20 and 22, causes the membranes 19 and 21 to tear open. The combustible fuel component and the oxidizer then flow into the combustion chambers 20 and 22 by virtue of the difference between the pressure of the gas confined in the portion 16,18 and the pressure in the combustion chambers 20 and 22.
As the fuel components flow out of the respective portions 12 and 14 and 16,18, the volume available to the confined gas increases and, as a consequence, its pressure decreases. According to the law governing the adiabatic expansion of gases, the initial pressure of the gas, that is, the pressure of the confined gas at the time that the fuel components begin to flow out of the storage tank 10, and the final pressure of the gas, that is, the pressure of the gas when the storage tank 10 has been substantially emptied, are in a ratio which is equal to the ratio of the total volume of the storage tank 10 to the initial volume of the confined gas. It will be understood that the quantity and the initial pressure of the confined gas are such as to permit the fuel components to be substantially entirely emptied from the storage tank 10.
The initial volume of the confined gas in the part 18 of the portion 16,18 may, for example, amount to about 20 to 50% of the total volume of the storage tank 10. The initial pressure of the confined gas may, for instance, lie between 20 and 40 bars and advantageously is equal to approximately 30 bars. However, the initial pressure of the confined gas is dependent upon the type of power plant with which the rocket is provided, that is, the initial pressure of the gas depends upon the size and construction of the power plant and upon the requirements imposed upon the power plant. With these considerations in mind, it is pointed out that the initial pressure of the confined gas may suitably be between about 5 and 100 bars, although the values for the initial pressure of the gas given here are not to be construed as limiting the invention in any manner.
Referring now to FIG. 2, the storage tank or container for the liquid fuel of a liquid-fueled rocket is here seen to include two sections 30 and 32 arranged in tandem. The section 30 has a portion 34 which is filled with a combustible fuel component to the level indicated at 35 and another portion 36 which is filled with a body of gas.
The section 32 has a portion 38 which is filled to the level indicated at 39 with an oxidizer, and the section 32 also has a portion 40 filled with a body of gas.
The propulsion means for the rocket is located below the section 32 and includes a combustion space or combustion chamber 48. The section 30 is connected with the combustion chamber 48 via a conduit 46. A valve 50 is provided in the conduit 46 as well as in the conduit connecting the section 32 with the combustion chamber 48. The valves 50 are normally closed until the rocket is to be started. It will be appreciated that the sections 30 and 32 define an enclosed space in which the fuel components and the gas are confined.
Again, the construction and arrangement of the combustion chamber, thevalves and the conduit system for conveying the fuel components from the storage tank to the combustion chamber do not form part of the invention per se and are not, therefore, illustrated in detail.
In operation of the embodiment shown in FIG. 2, the sections 30 and 32 are respectively filled with a combustible fuel component and an oxidizer to the respective levels 35 and 39, that is, the fuel components only partially fill the sections 30 and 32. The portions 36 and 40 of the sections 30 and 32, respectively, are then filled with a gas the quantity of which is sufficient in each instance to raise the pressure of the gas in each of the sections 30 and 32 to a value which exceeds the value of the pressure in the combustion chamber 48.
When the rocket is to be started or fired, the valves 50 are opened whereupon the fuel components are forced into the combustion chamber 48 by virtue of the difference between the pressure of the confined gas in the respective sections 30 and 32 and the pressure in the combustion chamber 48. The initial pressure of the gas in the section 30 may be different from that of the gas in the section 32 and, in the illustrated tandem construction, the initial pressure of the gas in the section 30 is advantageously lower than that of the gas in the section 32 in order to provide compensation for the greater hydrostatic pressure head of the fuel component in the section 30 as opposed to the lower hydrostatic pressure head of the fuel component in the section 32, i.e. in order to provide an equalization of pressure. (The hydrostatic pressure head in the section 32 falls off to substantially zero which, with regard to the section 30, is not the case because of the presence of the conduit 46 which latter is usually not completely emptied). It will be understood that the quantity and initial pressure of the gas in the respective sections 30 and 32 are such as to permit the fuel components to be substantially completely emptied therefrom.
In many circumstances, the sections 30 and 32 may be connected by a conduit 42 so as to permit an equalization of the pressures in the sections 30 and 32 through the conduit 42. Advantageously, the conduit 42 is provided with a valve 44.
If desired, the initial volume of the gas in the section 30 and the initial volume of the gas in the section 32 may be so selected that the ratio of the initial volume of the gas in the section 30 to the total volume of the section 30 equals the ratio of the initial volume of the gas in the section 32 to the total volume of the section 32.
By utilizing the invention, the total energy required for conveying the fuel components to the combustion chamber or chambers is already stored, in form of the compression of the gas, before the rocket is started or fired and, therefore, this energy need not be supplied to the storage tank or tanks from an external source. This manner of conveying the fuel components, that is, conveying the fuel components without introducing a gas or supplying energy to the storage tanks from externally, is structurally simpler and also cheaper than the conveying methods used for rockets heretofore. Furthermore, the entire operation is considerably more reliable when using the invention since there is no need to provide active regulating elements such as, for example, pressure reducing valves or gas generators. Such reliability is, of course, of great importance in rockets.
ROCKET
ENGINE
US3640072
US3640072
A rocket engine has an internal combustion chamber provided with a front wall. An outlet nozzle is provided in the front wall. At least two injection conduits communicate with the chamber rearwardly of the front wall in such a manner as to inject into the chamber respective streams of reactive propellants in direction tangentially of the chamber walls thus providing a short heat conduction path from the nozzle throat to the injected but yet unburned propellants rotating at high speed along the chamber walls.
BACKGROUND OF THE INVENTION
The present invention relates generally to a fuel-combusting device, and more particularly to a rocket engine. The invention also relates to a method of operating a rocket engine.
Rocket engines, and the operation thereof, are well known. The present invention is particularly concerned with small rocket engines wherein two or more liquid and/or gaseous fuel are injected for producing a gas stream. Such rocket engines are employed where small or very small amounts of thrust are needed, for instance as control thrusters of satellites, rocket-propelled aerospace vehicles and guided missiles. They are also used as the basic components of gas generators producing working gases such as are needed for the drive of turbines of auxiliary aggregates.
Rocket engines for these general purposes are of course already known. However, they suffer from various disadvantages, relating primarily to the problem of cooling the engines, providing proper propellant mixture ratio in the engine and operating the engine continuously. Particularly where small propellant quantities and small or very small thrusts below 7 pounds are involved, no properly operational rocket engines based on two or more component propellant systems are known, because it has been impossible to solve the cooling problem involved.
SUMMARY OF THE INVENTION
It is, accordingly, an object of the present invention to avoid the aforementioned disadvantages.
More particularly it is an object of the present invention ro provide a propellant-combusting device for rocket engines of low thrust such as used for the control of satellites, rocket-propelled aerospace vehicles and guided missiles, and also of the type which is used for gas production in gas generators.
A more particular object of the present invention is to provide such a device which provides for proper cooling, particularly in the region of the outlet nozzle.
An additional object of the invention is to provide a method of operating such a device.
In pursuance of the above objects, and others which will become apparent hereafter, one feature of my invention resides, briefly stated, in a propellant-combusting device which comprises wall means surrounding and defining an internal combustion chamber having a front wall portion. Outlet nozzle means is provided in the front wall portion and communicates with the chamber. Injecting means also communicates with the chamber and is operative for injecting into the same streams of reactive propellants in direction tangentially of the chamber, whereby the injected propellants initially sweep over and cool the wall means rearwardly of the front wall portion by taking off all heat conducted from the nozzle through the front wall radially outward simultaneously undergoing intimate mixture, prior to advancing towards and into the outlet nozzle means.
Because the rotation of the fuel streams resulting from the centrifugal forces acting upon them, and the resulting sweep of the fuel streams over the walls of the combustion chamber rearwardly of the front wall portion which is provided with the outlet nozzle, serves to cool these walls the present invention overcomes the cooling problem associated with the constructions known from the prior art. The flow of the fuel through the combustion chamber is radially inwardly from the outside towards the centrally located outlet nozzle and the maximum heat density is in the region of the throat of the outlet nozzle. According to the present invention the total cross-sectional area of the front wall portion in which the outlet nozzle is provided is a multiple of the cross-sectional area of the smallest radius of the throat of the outlet nozzle. In such a construction the heat transmitted to the outlet nozzle by the escaping hot gases is initially transmitted to the front wall portion surrounding the outlet nozzle, then radially outwardly conducted in this front wall portion, and then conducted rearwardly into the wall surrounding the combustion chamber rearwardly of the front wall portion into the region of the injecting means which injects the fuel components into the combustion chamber. In the region of injecting means the thus-conducted heat transmitted through the wall bounding the internal combustion chamber to the fuel which has been injected and which sweeps in a rotary motion over the inner surface of the wall under the influence of centrifugal force. It thus preheats the fuel which is desirable but is conducted away from and unable to damage the nozzle throat.
The novel features which are considered as characteristic for the invention are set forth in particular in the appended claims. The invention itself, however, both as to its construction and its method of operation, together with additional objects and advantages thereof, will be best understood from the following description of specific embodiments when read in connection with the accompanying drawing.
BRIEF DESCRIPTION OF THE DRAWING
FIG. 1 is a somewhat diagrammatic axial section through a device according to the present invention in one embodiment;
FIG. 2a is a section taken on the line A--A of FIG. 1;
FIG. 2b is a section analogous to FIG. 2a of an embodiment utilizing three injection means instead of two as in FIGS. 1 and 2a;
FIG. 3 is a section taken on the line B--B of FIG. 1;
FIG. 4 is a view similar to FIG. 1 but showing a further embodiment of the invention;
FIG. 5 is a view similar to FIG. 4 but showing still another embodiment of the invention;
FIG. 6 is a view similar to FIG. 5 showing yet an additional embodiment of the invention;
FIG. 7 is a further axial sectional view through another embodiment of the invention;
FIG. 8 is a view analogous to FIG. 7 but showing still an additional embodiment of the invention;
FIG. 9 is another axial section through still a further embodiment of the invention;
FIG. 10 is an axial section through a gas generator embodying the invention;
FIG. 11 is a view similar to FIG. 9 showing still a further embodiment of the invention; and
FIG. 12 is a view similar to FIG. 11 but showing yet another embodiment of the invention.
DESCRIPTION OF THE PREFERRED EMBODIMENTS
Discussing firstly the embodiment illustrated in FIGS. 1, 2a and 3, it will be understood that reference character E identifies the engine in general which comprises an internal combustion chamber 1 bounded by a rear wall 2, a front wall 3 and a circumferential wall 4. The front wall 3 is provided with an outlet nozzle 5 which is illustrated diagrammatically and whose particular construction may be in accordance with the teachings of the prior art well known to those skilled in this field. However, the nozzle 5 is located centrally of the front wall 3 and, in accordance with the present invention, the smallest cross-sectional diameter 8 of the nozzle 5 at the throat or neck thereof, is considerably smaller than the cross-sectional area of the front wall 3. This could also be stated, conversely, by saying that the cross-sectional area of the front wall 3 is a multiple of the cross-sectional area of the throat 8 of the nozzle 5.
The circumferential wall 4 is provided in the embodiment of FIGS. 1, 2a and 3 with two oppositely located inlet bores 6 and 7 constituting injecting means for two reactive propellants or propellant components, and in accordance with the invention and as clearly visible in FIG. 2a, bores 6 and 7 are so located that the streams of fuel injected into the combustion chamber 1 are injected tangentially to the periphery of the combustion chamber 1. They are thus forced to sweep over the inner surface of the circumferential wall 4 bounding the combustion chamber 1 in a rotary motion and to cool the wall 4, before they mix and react with one another. The bores 6 and 7 are located in the embodiment of FIGS. 1, 2a and 3 in a common plane normal to the axis of the combustion chamber 1, that is the axis extending through the nozzle 5. In the embodiment illustrated in FIG. 2b, which corresponds to that of FIGS. 1, 2a and 3 in most particulars, there are provided three inlet bores 6, 7 and 15 which each inject a stream of a propellant component. It is evident from FIG. 2b that the three bores 6, 7 and 15 may also be located in a common transverse plane.
FIGS. 1 and 3 shown the manner in which heat is conducted away from the nozzle 5 in operation of the engine E. The heat transmitted to the nozzle 5 by the escaping hot gases is conducted radially through the front wall portion 3 (see FIG. 3) and is then conducted rearwardly into the circumferential wall 4 to the region of the inlet bores 6 and 7 (and 15, in the case of FIG. 2b) where it is transmitted to the injected propellants which sweeps over the inner surface of the circumferential wall 4 prior to mixture. Because of the nonlinear radial temperature curve in the wall portion 3 which has a very large cross-sectional area by comparison with the cross-sectional area of the throat 8 of the nozzle 5, this manner of conducting heat away from the nozzle 5 is the more advantageous the smaller the cross-sectional area of the throat 8 of the nozzle 5 is, that is the smaller the thrust of the engine or the smaller the flow of fuel therethrough. By contrast to what is known from the art, the novel construction provides a cooling effect which is not only achieved in a most simple manner but which is extremely reliable and which is afforded in particular for the throat 8 of the nozzle 5, that is that portion of the engine which is subjected to the most heating and therefore susceptible of the most damage by having the fuel come in contact therewith, although in FIG. 5 the lines indicating the fuel flow are not shown in actual contact with this inner surface 14a for the sake of clarity.
Coming now to the embodiment shown in FIG. 6 it will be seen that here the configuration of the combustion chamber 18 is the reverse of that in FIG. 5, that is that the combustion chamber diverges conically in direction towards the outlet nozzle, rather than away therefrom. In this embodiment, also, there are provided a plurality of inlet bores 6 for one fuel component and inlet bores 7 for the other fuel component, with one bore 6 and one bore 7 always being located in one common transverse plane 20 extending transversely of the elongation of the chamber 18. Of course, the inlet bores 6 and 7 need not all be located on one side, and the construction could be modified so that on one plane 20 the inlet bore 6 is located at the left-hand side and on the next plane 20 the inlet bore 6 is located at the right-hand side of the illustration in FIG. 6. In any case, however, each of the inlet bores 6 and 7 is controlled by a separate control valve 17 but each inlet bore may be separately opened and closed, to thereby vary the throughput of fuel and the thrust in simple and highly effective manner by adding or taking away the output of individual ones of the bores 6 and 7. This control arrangement is the one which has been suggested in connection with the embodiment in FIG. 4, and can of course be employed in that embodiment.
In FIG. 7 I have illustrated a construction wherein the internal combustion chamber 22 is of semicircular configuration. All other features are the same as in the preceding embodiments, and like reference numerals identify like components. The heat flow from the nozzle is the same as identified in FIG. 3, and this is true in all embodiments already described and those still to be discussed. Also, the injection of the fuel fluids is always tangential, both in the embodiments which have been discussed here before and in those which are still to be described.
FIG. 8 shows a construction wherein the combustion chamber 24 is of spherical configuration and wherein the front wall portion 3 can be considered to extend from the region of the inlet bore 6 to the region of the inlet bore 7.
In the embodiment shown in FIG. 9 the combustion chamber 26a is composed of a series of substantially barrel-shaped sections 26, with the injection of fuel fluids taking place through the conduits or inlet bores 6 and 7 on three different transverse planes 28--each corresponding to one of the barrel-shaped sections 26 and bisecting the same at its greatest diameter. Each of the inlet bores 6 and 7 can of course be separately controlled in the same manner as discussed with respect to FIG. 6.
FIG. 10 shows a gas generator embodying the present invention. It comprises an internal combustion chamber 1 provided with the inlet bores 6 and 7 and corresponding to the embodiment illustrated in FIG. 1. Reference numeral 32 identifies the outlet nozzle whose downstream or outlet end communicates with a second internal combustion chamber 30 which again is provided with inlet bores 11 and 13 for two fuel fluids, both of which are also injected tangentially in the same manner as takes place in the chamber 1. As mentioned within the discussion of the embodiment in FIG. 1, the throat 8 of the outlet nozzle 32 is again so dimensioned that its cross-sectional area is much smaller than the cross-sectional area of the front wall portion 3 separating the chambers 1 and 30 from one another, with the flow of conducted heat being illustrated by the arrows in FIG. 9, from which it will be seen that the heat is here transmitted from the front wall portion 3 towards the inlet bores 6 and 7 as well as towards the inlet bores 11 and 13. The gases issuing through the nozzle 32 into the chamber 30 encounter additional fuel injected through the inlet bores 11 and 13 in a nonstoichiometric relationship, so that the hot gases issuing through the nozzle 32 into the chamber 30 are strongly cooled. In place of additional fuel, or in addition to such additional fuel, it is also possible to inject other means serving to cool the hot gases entering the chamber 30 from the chamber 1. However, the injection should always take place in the vicinity of the front wall portion 3 because this provides for additional cooling.
Finally, the embodiment illustrated in FIG. 12 shows an engine according to the present invention wherein the internal combustion chamber 38 is of substantially lenticular configuration and wherein the nozzle 40 is of the type known as a corner expansion nozzle. The injection of the propellants through the inlet bores 6 and 7 is of course again tangentially to the circumference of the chamber 38.
ROCKET
DRIVE COOLING ARRANGEMENT
US3462956 // GB1196489
US3462956 // GB1196489
Improvements in or relating to Rocket Engines I, LUTZ TILO KAYSER, a Citizen of the Federal Republic of Germany, of Am Bismarckturm 10, Germany, do hereby declare the invention, for which I pray that a patent may be granted to me, and the method by which it is to be performed, to be particularly described in and by the following statement: This invention relates to rocket engines and provides a method of cooling rocket engine combustion chambers for liquid or gaseous propellents and in which mixtures of these two types of propellents may be used with the same advantage.
For rocket engines, several types of cooling methods and devices have already been developed, but they can in no way be considered as perfect. The following four main types of cooling systems are concerned:
1. The so-called regenerative cooling in which the propellent is made, before injection, to absorb the heat transferred to a cooling jacket completely enclosing the nozzle and the combustion chamber. This process, however, is too complicated for small rocket engines and represents an explosion hazard when applied to temperature sensitive propellents.
2. In the case of so-called ablative cooling, the entire combustion chamber and the nozzle are formed of a material which is a poor heat conductor. Also a carbon or a preferably fibrereinforced synthetic resin forming a viscous melt may be used. The surface of such material assumes the combustion gas temperature when the device is in operation, and the heat is radiated back to the combustion chamber.
A combustion chamber constructed on this basis, however, involves the risk of asymmetrical burning at the nozzle due to fluctuating wear phenomena.
3. In so-called film and transpiration cooling, a cooling effect of the parts of the wall to be protected is produced by introduction of propellents through pores or fine holes along the walls of the combustion chamber and the nozzle. Due to the expulsion of unconsumed propellent in the boundary layers of the flow, a reduction of performance occurs.
4. With the use of so-called radiation cooling, the combustion chamber and nozzles must be manufactured from materials having a high melting point, since they almost reach the combustion temperature and radiate heat to the environment, the surrounding parts of the apparatus also being heated.
The above-stated and also other disadvantages are the reasons why improved cooling measures and systems for rocket engines are still required.
It is a particular problem that in a rocket engine the maximum heat transferred from the combustion gases to the wall of the combustion chamber occurs at the throat and a small range upstream and downstream of the narrowest cross-section thereof.
It is an object of the present invention to avoid or reduce the above-stated defects and disadvantages of various known cooling systems for rocket engines.
Accordingly, the present invention provides a rocket engine in which at least one of the propellent components, for example the fuel or the oxidizer, is injected onto an outwardly facing concavely curved surface of a toroidal cooling chamber formed around the nozzle throat, and from there flows to the inside surface of the combustion chamber wall to flow along said wall in counterflow to the hot combustion gases, with a film cooling effect.
Both the fuel and the oxidizer, or either one of them, may be fed into the rocket engine through radially inwardly directed injection apertures, or through nozzles.
Both propellents combined, or only one of them may be first radially injected from outside on to the outwardly facing concavely curved surface of the toroidal cooling chamber formed around the nozzle throat. This achieves intensive cooling of this part of the nozzle which is most intensely heated from inside.
Regenerative cooling, in the conventional meaning of passing the coolant through a narrow gap between two walls, one of which is to be cooled, is not used in the present case therefore.
The propellent then flows from the outside of the nozzle, with an approximately radiall3 outward flow direction and is forced against the concave inner wall of the combustion chamber also by centrifugal force, in order to flow along this wall in counterflow to the hot combustion gases, with a cooling film effect, Combustion begins here, the combustion gases flowing to the centre of the chamber and the nozzle and the unburned liquids continuing to flow along the wall. A combustion chamber of a rocket engine, provided with a cooling system according to the present invention, can be constructed as a cone, as a cylinder or as a ring.
Experimental tests have shown that, with higher thrust rocket engines, it is an advantage only to inject one component of the propellents at the nozzle and to inject the other component into the combustion chamber. The feeding of both propellents combined at the nozzle throat of a system of comparatively large dimensions, results in excessively long mixing paths and consequently in premature reaction in the vicinity of the nozzle throat, that is to say, it results in a reduced cooling action.
The other propellent component is then fed in counterflow to the combustion gases of the combustion chamber tangentially to the inside of the wall of the combustion chamber.
In another variation the propellents can be injected with a tangential velocity component in addition to their radial velocity components in order to further improve the mixing and combustion.
In order that the invention may readily be carried into practice, embodiments thereof will now be described in detail, by way of example, with reference to the accompanying drawings, in which:
Figure 1 is an axial longitudinal section through a single chamber of a rocket engine fitted with a cooling system according to the invention;
Figure 2 is an axial longitudinal section through a rocket engine similar to Figure 1, several possibilities of feeding the propellents being shown;
Figure 3 is a section through an annular rocket engine which is obtained geometrically by rotation of the longitudinal section according to Figure 1 through an eccentric axis of Figure 1, the cooling arrangement being constructed similarly as in Figure 1; and
Figure 4 is another embodiment of an annular rocket engine.
Figure 1 represents the embodiment of a rocket engine. The fuel is fed, in this example, through a duct 1 and the oxidizer through a duct 2, by way of respective annular propellent passages 3 and 4. The two 1 propellents are injected along the sides of an annular diaphragm 6, flowing in a radial direction into a cooling chamber 7, formed as a toroidal cavity around the throat 10 of a nozzle 8. Due to the high injection speed and the resultant centrifugal force, the liquid and/or gas flow is pressed powerfully against an outwardly facing concavely curved surface of the toroidal chamber 7, as shown in broken lines in Figure 1.
Intensive regenerative cooling is achieved at this part of the nozzle, which is most intensely heated at its throat 10.
The cooling action is not achieved by a cooling jacket separating the propellents from each other, but the cooling action begins only after the injection of the propellents, that is to say, at combustion chamber pressure. Thus the action is not regenerative cooling, in the conventional manner and the whole injection velocity is available for cooling.
The mixed propellents now flow, under the effect of centrifugal force, along the inside of the combustion chamber wall 9, in counterflow to the combustion gases into the combustion chamber 12, with a cooling film effect.
They react there with each other and so form the combustion gases, which flow out of the nozzle 11. As already mentioned, with propellents having a very high reaction speed or with rocket engines having higher thrust, it is an advantage to inject only one of the two components, in the manner described and to feed the other component tangentially to the inside of the combustion chamber wall 9.
Figure 2 shows several alternatives of separate propellent injection. To avoid ambiguity, it is here emphasized that several manners of feeding propellents are illustrated with reference to this drawing, only a specific combination being illustrated in detail.
In detail the following possibilities of combination are achieved: a) One propellent component, for example the fuel, is injected radially from outside through a duct 13 onto the outwardly facing concavely curved surface of the toroidal cooling chamber, and is then diverted towards the inner wall of the combustion chamber, b) One component, as described under (a) is injected through the duct 13, whilst the other component is fed through a duct 15, tangentially to the inner surface of the combustion 120 chamber wall on the lower side of the combustion chamber.
c) One component is injected through a duct 16, more or less inclined onto the outwardly facing concavely curved surface of the toroidal cooling chamber, whilst the other component is fed through the duct 15 of the combustion chamber.
d) One component is fed through the duct 16 and the other component is fed through the duct 14. According to the type of application, the duct 14 can be displaced more towards the top of the combustion chamber or more towards the underside thereof.
Due to the above-mentioned alternatives of place of the propellent injection, the cooling arrangement according to the present invention can be adapted to the thrust of the rocket engine and the reaction speed of the propellents employed.
In Figure 4, the propellent or propellent components is or are supplied through the ducts 17 and 18.