Exploring Advanced Rocket Engine Technology
Rocket Engines are basically used to store the rocket mass of rocket propellant for forming its high speed propulsive jet. Those vehicles which are propelled by rocket engines are commonly known as rockets. Rockets function on the principle of Newton’s 3rd law of motion. We always see that most of the rocket engines use combustion but non-combusting such as cold gas thrusters forms the also exist. In comparison, of rocket engine to other types of jet engines, rocket engines are by far the lightest, have the highest thrust and have least propellant efficiency. Rocket engines become more efficient with higher velocities, due to greater propulsive efficiency and the Obvert effect. Since they do not require any atmosphere, they are well suited for uses at very high altitudes and in space.
Today, human culture or scientists have discovered the new type of rocket engine which consumes less amount of money and having high-efficiency comparing to the start up by the formation of the Rocket engine. Hence in the future, there is a lot of amount of Innovation which will happen in the field of the rocket engine. Today’s scientist was getting work on the new type of rocket engine which gets work on the phenomena plasma. Today NASA and some other organization get work to send the rocket to Mars which required a heavy amount of fuel to getting cover the distance up to mars. Which increase the mass of rocket so, scientists are getting to work on different type of rocket engines such as plasma type.
Rocket Engine applications
Typically the propulsion systems and thus rocket engines are tailored to the specific requirements of launching the concept. This is has consequences insofar as any improvement or modification of the engine has significant impact on the entire launch system and even more important an improvement of the engines performance may not necessarily improve overall efficiency of the launcher. The currently operational space transportation systems can be divided into the two major types Launchers with large strap-on boosters with the core and upper stages such as the European ARIANE Japanese HII the Russian Soyuz or the American Space Shuttle and launchers with the booster an optional sustained and an upper stage such as the American ATLAS and Delta the Russian PROTON the Chinese Long March or the international Sea Launch the Vehicle.
History of Rocket Engine
According to the ancient Roman writings, AulusGellius Inc. 400 BC a Greek Pythagorean named Archaism propelled the wooden bird along wires using the steam. However it is a would not appear because it did not have so much amount of the thrust. In the 13th century a turning point in the rocket technology emerged with a short typescript the entitled LiberIgniumadComburendos Hosts abbreviated as The Book of Fires. The typescript is composed of recipes for the creating incendiary weapons from mid-eighth to the end of thirteenth century two of which are the rockets.
The first recipe calls for one part, which is Sulphur, added to six parts of saltpeter (potassium nitrate) dissolved in laurel oil, then inserted into hollow wood and lit to actually fly away whenever you want, to whatever destination you wish and burn up everything. The second recipe combines a pound of Sulphur, two pounds of charcoal, and six pounds of saltpeter all finely powdered on a marble slab. Then this powder mixture is the packed firmly into a long and narrow the case. The saltpeter’s introduction into pyrotechnic mixtures connected the shift from hurled Greek fire into self-propelled rocketry.
Components of Liquid Rocket Engines
The key components of a liquid propellant rocket engine are the devices for propellant delivery the turbines and pumps as well as the generators of the driving gases gas generator or pre-burner the propellant injection system the thrust chamber which combines the combustion chamber and a short part of the diverging section of the nozzle which typically ends with the distribution manifold for propellants used as coolant for the thrust chamber liner and finally the thrust nozzle. Each of these devices has it’s the own design problems and criteria and the optimum of one component not necessarily leads to optimum of the entire system. The following section will be touch briefly the key issues of each of the components but will finally focus on the thrust chamber.
1.Gas generator and perjurer
The main requirement for any gas generator or perjurer is the delivery of a sufficient amount of the driver gas a designed pressure and temperature which the guarantees a continuous propellant supply of the thrust chamber. A quick glance at the equation (4), yields that the necessary turbine power depends aside of the desired ratio of exit to entry pressure and the mass flow rate of gases solely on thermodynamic properties the temperature heat capacity and isentropic the coefficient and of course efficiency of the device.
As the already mentioned previously turbine exit pressure in the gas generator cycle engine is the independent of pressure in the thrust chamber. This is an entirely different for the staged combustion engines where the exhaust gases are fed to the chamber. While the combustion pressure in the gas generator lay frequently below one in the thrust chamber pressures in the pre-burners are typically a factor of two or the three higher see table 6 which summarizes characteristic data of the sample gas generators and prate burners. A high turbine power may be the achieved by appropriate mass flow rates and the pre-burner exhaust gas the temperatures.
The function of turbo pump system of the rocket engine is to receive the liquid propellants from the vehicle tanks at low pressure and supply them to the combustion chamber at required flow rate and the injection pressure. The energy to power the turbine itself is provided by the expansion of high pressure gases which are the usually mixtures of propellants being the pumped. Radial pumps and axial turbines are of the common use. Figure 5 basic elements of a turbo pump and their main function as well as the flow passages.
3.Thrust chamber assembly
In a liquid propellant rocket engine the function of thrust chamber assembly is to the generate thrust by efficiently converting the propellant chemical energy into hot gas kinetic the energy. That conversion is accomplished by the combustion of propellants in the combustion chamber followed by acceleration of the hot gas through a convergent divergent nozzle to the achieve high gas velocities and the thrust.
The basic elements of the thrust chamber assembly of a rocket engine consists of the following components which are closely coupled in their functioning and the therefore have to be designed in parallel the card an welded to the injector head which is which holds the propellant distribution manifolds and the ignition system and the combustion chamber liner including the throat section and first part of the nozzle to which the inlet manifold for coolant is the welded. Looking at the figure 9 a sketch of the Vulcan thrust chamber simplifies the explanation of these parts.
Advanced design methodology
A clear alternative to the deterministic approach briefly described in the chapter 2.3.5 is a probabilistic design approach which considers the uncertainties of material properties operating conditions and the loads in a more structured the manner. In the probabilistic approach design variables are the neither seen as single values nor are they weighed to an upper or the lower bounds. They are instead represented by the actual distribution are variation of the parameters. A distribution is the histogram of discrete values of the parameters or a mathematical model that represents a smooth description of the variation.
When the area under the histogram or curve is normalized to a value of one the function is called the probability density the function. The parameters are termed random the variables. The distribution functions then are used to the determine probability of occurrence of the given value of random the variable. A typical structural model for both static and the dynamic analyses bases on nominal hardware the geometry. Resulting displacements loads and the stresses are then compared to the material data cyclic fatigue or ultimate the strength.
The material data are taken as lower bound values of the available material test the data. The ratios between the data and the predicted values have to be within specified safety the margin. The predicted distributions show the probability for a failure and not a value or the safety. The sensitivity of each variable to failure which is the part of result and this information helps the designer to get a better feeling for impact of the variable. It is a essential to the become knowledgeable of inherent risk of the failure and assess it identify the mayor causes and minimize them within design the constraints.
The new concept or design has to show the capability to improve reliability performance or cost effectiveness of the system or component and in the recent years reliability and cost became the dominating. Hence current developments deal with the low cost approaches such as the RS-68 or more recent European Vulcan X design the studies. Within this chapter a series of the new concepts and a novel design methodology approach are the discussed and their potential contribution to the above mentioned drivers will be the show.
1.Advanced ignition devices
An ignition system is the designed to meet requirements of the mission sea level vacuum or multiple ignitions and the accounts for further boundary conditions such as the propellant phase and temperature as well as geometric and the power constraints. Any improvement will be the have to account for these conditions and will be the most likely be mission specific as the well.
The simplest form of system improvement is the reduction of complexity by reducing the number of parts. This is the may be combined with a reduction of component the weight. An example could be to use the original propellants instead of the additional ones. For the re-ignitable engines a simple pyrotechnic system isn’t sufficient and another form has to be the used. The Quite often gaseous pressurized propellants are applied which are send to and the ignition chamber at the appropriate mixture ratio and then injected into the combustion chamber.
Multiple ignitions require a sufficient amount of the propellant and thus large enough the tanks. The entire system including valves ignition chamber also may be the omitted if another system with similar reliability becomes the available. Although not yet used in cars yet the automotive industry currently develops igniters for their internal combustion engines which the apply semiconductor lasers as we already use in the our MP3 or CD players. A change from the continuous wave to pulse mode operation allows for a considerable increase in the optical power.
Advanced injection technique
High precision fabrication qualification inspection and the testing high precision mounting and the quality assurance procedures for injector are time consuming and the costly. There are the two possible ways to alter the procedures mentioned above first a reduction of the precision requirements through the simpler injector design or second a reduction of the number of injection elements the through an increase in injector mass flow the rate.
Based on a description of the basic components and features of the rocket engines areas have been identified and the physical and technological limitations given in the order to pave way for a discussion of possible the improvements. The topics of gas generator or pre-burner as well as were not included in the areas where potential advancements were proposed the although some of techniques proposed for the thrust chambers ignition systems or injector should be the applicable as well for other combustion the devices. The concept of an advanced the laser-based ignition system was discussed and current status the given.
Additionally an the injection system design which aims at a much cheaper the design with negligible impact on performance with the simple LOX posts and a porous face plate was proposed and the sample results were given. Various possibilities to the improve thrust chamber cooling techniques involving new the materials such as Corn ceramic matrix composites or the effusion cooling were explained the different aspects shown. Additionally the advantages and basic features of and the advanced nozzle concept for a main stage engine the dual bell nozzle was described and the current statues of technology the explained. Finally a novel design methodology the probabilistic design approach was briefly the described.