PDF | Starting with some basics about space transportation systems Rocket engines may either work with solid or liquid propellants or as a. Program: to design and fly a human-rated reusable liquid propulsion rocket engine to launch the shuttle. It was the first and only liquid-fueled rocket engine to be. In rocket propulsion, a mass of propellant (m) is accelerated (via the http://ntrs medical-site.info
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A rocket engine uses stored rocket propellant mass for forming its high-speed propulsive jet. "Space Shuttle Main Engine" (PDF). Pratt & Whitney Rocketdyne. A rocket engine is not like a conventional engine. A conventional engine ignites fuel which then pushes on some pistons, and it turns a crank. Therefore, it uses. Types of Rocket Propulsion. • Solid. – Fuel and oxidizer coexist in a solid matrix. • Liquid. – Fluid (liquid or gas) propellants stored separately. – Propellants.
Unlike in airbreathing jet engines , no atmospheric nitrogen is present to dilute and cool the combustion, and the propellant mixture can reach true stoichiometric ratios. This, in combination with the high pressures, means that the rate of heat conduction through the walls is very high. In order for fuel and oxidizer to flow into the chamber, the pressure of the propellant fluids entering the combustion chamber must exceed the pressure inside the combustion chamber itself.
This may be accomplished by a variety of design approaches including turbopumps or, in simpler engines, via sufficient tank pressure to advance fluid flow. Tank pressure may be maintained by several means, including a high-pressure helium pressurization system common to many large rocket engines or, in some newer rocket systems, by a bleed-off of high-pressure gas from the engine cycle to autogenously pressurize the propellant tanks   For example, the self-pressurization gas system of the BFR is a critical part of SpaceX strategy to reduce launch vehicle fluids from five in their legacy Falcon 9 vehicle family to just two in BFR, eliminating not only the helium tank pressurant but all hypergolic propellants as well as nitrogen for cold-gas reaction-control thrusters.
The hot gas produced in the combustion chamber is permitted to escape through an opening the "throat" , and then through a diverging expansion section. When sufficient pressure is provided to the nozzle about 2. Exhaust speeds vary, depending on the expansion ratio the nozzle is designed for, but exhaust speeds as high as ten times the speed of sound in air at sea level are not uncommon.
About half of the rocket engine's thrust comes from the unbalanced pressures inside the combustion chamber, and the rest comes from the pressures acting against the inside of the nozzle see diagram. As the gas expands adiabatically the pressure against the nozzle's walls forces the rocket engine in one direction while accelerating the gas in the other.
The most commonly used nozzle is the de Laval nozzle , a fixed geometry nozzle with a high expansion-ratio. The large bell- or cone-shaped nozzle extension beyond the throat gives the rocket engine its characteristic shape. The exit static pressure of the exhaust jet depends on the chamber pressure and the ratio of exit to throat area of the nozzle. As exit pressure varies from the ambient atmospheric pressure, a choked nozzle is said to be.
In practice, perfect expansion is only achievable with a variable-exit area nozzle since ambient pressure decreases as altitude increases , and is not possible above a certain altitude as ambient pressure approaches zero.
If the nozzle is not perfectly expanded, then loss of efficiency occurs. Grossly over-expanded nozzles lose less efficiency, but can cause mechanical problems with the nozzle.
Fixed-area nozzles become progressively more under-expanded as they gain altitude. Almost all de Laval nozzles will be momentarily grossly over-expanded during startup in an atmosphere. Nozzle efficiency is affected by operation in the atmosphere because atmospheric pressure changes with altitude; but due to the supersonic speeds of the gas exiting from a rocket engine, the pressure of the jet may be either below or above ambient, and equilibrium between the two is not reached at all altitudes see diagram.
For optimal performance, the pressure of the gas at the end of the nozzle should just equal the ambient pressure: To maintain this ideal of equality between the exhaust's exit pressure and the ambient pressure, the diameter of the nozzle would need to increase with altitude, giving the pressure a longer nozzle to act on and reducing the exit pressure and temperature. This increase is difficult to arrange in a lightweight fashion, although is routinely done with other forms of jet engines.
In rocketry a lightweight compromise nozzle is generally used and some reduction in atmospheric performance occurs when used at other than the 'design altitude' or when throttled. To improve on this, various exotic nozzle designs such as the plug nozzle , stepped nozzles , the expanding nozzle and the aerospike have been proposed, each providing some way to adapt to changing ambient air pressure and each allowing the gas to expand further against the nozzle, giving extra thrust at higher altitudes.
When exhausting into a sufficiently low ambient pressure vacuum several issues arise. One is the sheer weight of the nozzle—beyond a certain point, for a particular vehicle, the extra weight of the nozzle outweighs any performance gained. Secondly, as the exhaust gases adiabatically expand within the nozzle they cool, and eventually some of the chemicals can freeze, producing 'snow' within the jet.
This causes instabilities in the jet and must be avoided. On a de Laval nozzle , exhaust gas flow detachment will occur in a grossly over-expanded nozzle. As the detachment point will not be uniform around the axis of the engine, a side force may be imparted to the engine. This side force may change over time and result in control problems with the launch vehicle. Advanced altitude-compensating designs, such as the aerospike or plug nozzle , attempt to minimize performance losses by adjusting to varying expansion ratio caused by changing altitude.
For a rocket engine to be propellant efficient, it is important that the maximum pressures possible be created on the walls of the chamber and nozzle by a specific amount of propellant; as this is the source of the thrust.
This can be achieved by all of:. Since all of these things minimise the mass of the propellant used, and since pressure is proportional to the mass of propellant present to be accelerated as it pushes on the engine, and since from Newton's third law the pressure that acts on the engine also reciprocally acts on the propellant, it turns out that for any given engine, the speed that the propellant leaves the chamber is unaffected by the chamber pressure although the thrust is proportional.
However, speed is significantly affected by all three of the above factors and the exhaust speed is an excellent measure of the engine propellant efficiency. This is termed exhaust velocity , and after allowance is made for factors that can reduce it, the effective exhaust velocity is one of the most important parameters of a rocket engine although weight, cost, ease of manufacture etc.
For aerodynamic reasons the flow goes sonic " chokes " at the narrowest part of the nozzle, the 'throat'. Since the speed of sound in gases increases with the square root of temperature, the use of hot exhaust gas greatly improves performance.
Expansion in the rocket nozzle then further multiplies the speed, typically between 1. The speed increase of a rocket nozzle is mostly determined by its area expansion ratio—the ratio of the area of the throat to the area at the exit, but detailed properties of the gas are also important.
Larger ratio nozzles are more massive but are able to extract more heat from the combustion gases, increasing the exhaust velocity. Vehicles typically require the overall thrust to change direction over the length of the burn. A number of different ways to achieve this have been flown:. Rockets can be further optimised to even more extreme performance along one or more of these axes at the expense of the others.
An engine that gives a large specific impulse is normally highly desirable.
The specific impulse that can be achieved is primarily a function of the propellant mix and ultimately would limit the specific impulse , but practical limits on chamber pressures and the nozzle expansion ratios reduce the performance that can be achieved. Below is an approximate equation for calculating the net thrust of a rocket engine: Since, unlike a jet engine, a conventional rocket motor lacks an air intake, there is no 'ram drag' to deduct from the gross thrust. Consequently, the net thrust of a rocket motor is equal to the gross thrust apart from static back pressure.
At full throttle, the net thrust of a rocket motor improves slightly with increasing altitude, because as atmospheric pressure decreases with altitude, the pressure thrust term increases.
This reduction drops roughly exponentially to zero with increasing altitude. Maximum efficiency for a rocket engine is achieved by maximising the momentum contribution of the equation without incurring penalties from over expanding the exhaust. Since ambient pressure changes with altitude, most rocket engines spend very little time operating at peak efficiency. Due to the specific impulse varying with pressure, a quantity that is easy to compare and calculate with is useful.
It is thus quite usual to rearrange the above equation slightly: In liquid and hybrid rockets, the propellant flow entering the chamber is controlled using valves, in solid rockets it is controlled by changing the area of propellant that is burning and this can be designed into the propellant grain and hence cannot be controlled in real-time.
Rockets can usually be throttled down to an exit pressure of about one-third of ambient pressure  often limited by flow separation in nozzles and up to a maximum limit determined only by the mechanical strength of the engine. In practice, the degree to which rockets can be throttled varies greatly, but most rockets can be throttled by a factor of 2 without great difficulty;  the typical limitation is combustion stability, as for example, injectors need a minimum pressure to avoid triggering damaging oscillations chugging or combustion instabilities ; but injectors can be optimised and tested for wider ranges.
For example, some more recent liquid-propellant engine designs that have been optimised for greater throttling capability BE-3 , Raptor can be throttled to as low as 18—20 percent of rated thrust. Rocket engine nozzles are surprisingly efficient heat engines for generating a high speed jet, as a consequence of the high combustion temperature and high compression ratio.
Rocket nozzles give an excellent approximation to adiabatic expansion which is a reversible process, and hence they give efficiencies which are very close to that of the Carnot cycle. See Rocket energy efficiency for more details. Rockets, of all the jet engines, indeed of essentially all engines, have the highest thrust to weight ratio. This is especially true for liquid rocket engines. This high performance is due to the small volume of pressure vessels that make up the engine—the pumps, pipes and combustion chambers involved.
The lack of inlet duct and the use of dense liquid propellant allows the pressurisation system to be small and lightweight, whereas duct engines have to deal with air which has a density about one thousand times lower. Of the liquid propellants used, density is worst for liquid hydrogen. Although this propellant is marvellous in many ways, it has a very low density, about one fourteenth that of water. This makes the turbopumps and pipework larger and heavier, and this is reflected in the thrust-to-weight ratio of engines that use it for example the SSME compared to those that do not NK Most other jet engines have gas turbines in the hot exhaust.
Due to their larger surface area, they are harder to cool and hence there is a need to run the combustion processes at much lower temperatures, losing efficiency. Two exceptions are graphite and tungsten , although both are subject to oxidation if not protected. Indeed, many construction materials can make perfectly acceptable propellants in their own right.
It is important that these materials be prevented from combusting, melting or vaporising to the point of failure. This is sometimes somewhat facetiously termed an "engine-rich exhaust". Materials technology could potentially place an upper limit on the exhaust temperature of chemical rockets. Alternatively, rockets may use more common construction materials such as aluminium, steel, nickel or copper alloys and employ cooling systems that prevent the construction material itself becoming too hot.
Regenerative cooling , where the propellant is passed through tubes around the combustion chamber or nozzle, and other techniques, such as curtain cooling or film cooling, are employed to give longer nozzle and chamber life.
These techniques ensure that a gaseous thermal boundary layer touching the material is kept below the temperature which would cause the material to catastrophically fail.
The strongest heat fluxes are found at the throat, which often sees twice that found in the associated chamber and nozzle. This is due to the combination of high speeds which gives a very thin boundary layer , and although lower than the chamber, the high temperatures seen there. See rocket nozzles above for temperatures in nozzle. In all cases the cooling effect that prevents the wall from being destroyed is caused by a thin layer of insulating fluid a boundary layer that is in contact with the walls that is far cooler than the combustion temperature.
Provided this boundary layer is intact the wall will not be damaged. Disruption of the boundary layer may occur during cooling failures or combustion instabilities, and wall failure typically occurs soon after.
With regenerative cooling a second boundary layer is found in the coolant channels around the chamber. This boundary layer thickness needs to be as small as possible, since the boundary layer acts as an insulator between the wall and the coolant. This may be achieved by making the coolant velocity in the channels as high as possible.
Liquid-fuelled engines are often run fuel-rich , which lowers combustion temperatures. This reduces heat loads on the engine and allows lower cost materials and a simplified cooling system. This can also increase performance by lowering the average molecular weight of the exhaust and increasing the efficiency with which combustion heat is converted to kinetic exhaust energy.
When operated within significant atmospheric pressure, higher combustion chamber pressures give better performance by permitting a larger and more efficient nozzle to be fitted without it being grossly overexpanded.
However, these high pressures cause the outermost part of the chamber to be under very large hoop stresses — rocket engines are pressure vessels. Worse, due to the high temperatures created in rocket engines the materials used tend to have a significantly lowered working tensile strength.
In addition, significant temperature gradients are set up in the walls of the chamber and nozzle, these cause differential expansion of the inner liner that create internal stresses. The extreme vibration and acoustic environment inside a rocket motor commonly result in peak stresses well above mean values, especially in the presence of organ pipe -like resonances and gas turbulence.
The combustion may display undesired instabilities, of sudden or periodic nature. The pressure in the injection chamber may increase until the propellant flow through the injector plate decreases; a moment later the pressure drops and the flow increases, injecting more propellant in the combustion chamber which burns a moment later, and again increases the chamber pressure, repeating the cycle.
This may lead to high-amplitude pressure oscillations, often in ultrasonic range, which may damage the motor. The other failure mode is a deflagration to detonation transition ; the supersonic pressure wave formed in the combustion chamber may destroy the engine. Combustion instability was also a problem during Atlas development.
The Rocketdyne engines used in the Atlas family were found to suffer from this effect in several static firing tests, and three missile launches exploded on the pad due to rough combustion in the booster engines. In most cases, it occurred while attempting to start the engines with a "dry start" method whereby the igniter mechanism would be activated prior to propellant injection. During the process of man-rating Atlas for Project Mercury , solving combustion instability was a high priority, and the final two Mercury flights sported an upgraded propulsion system with baffled injectors and a hypergolic igniter.
The problem affecting Atlas vehicles was mainly the so-called "racetrack" phenomenon, where burning propellant would swirl around in a circle at faster and faster speeds, eventually producing vibration strong enough to rupture the engine, leading to complete destruction of the rocket. It was eventually solved by adding several baffles around the injector face to break up swirling propellant. More significantly, combustion instability was a problem with the Saturn F-1 engines.
Some of the early units tested exploded during static firing, which led to the addition of injector baffles.
In the Soviet space program, combustion instability also proved a problem on some rocket engines, including the RD engine used in the R-7 family and the RD used in the R family, and several failures of these vehicles occurred before the problem was solved. The combustion instabilities can be provoked by remains of cleaning solvents in the engine e. In stable engine designs the oscillations are quickly suppressed; in unstable designs they persist for prolonged periods.
Oscillation suppressors are commonly used. Periodic variations of thrust, caused by combustion instability or longitudinal vibrations of structures between the tanks and the engines which modulate the propellant flow, are known as " pogo oscillations " or "pogo", named after the pogo stick.
This is a low frequency oscillation at a few Hertz in chamber pressure usually caused by pressure variations in feed lines due to variations in acceleration of the vehicle. Chugging can be minimised by using gas-filled damping tubes on feed lines of high density propellants.
This can be caused due to insufficient pressure drop across the injectors. However, in extreme cases combustion can end up being forced backwards through the injectors — this can cause explosions with monopropellants. This is the most immediately damaging, and the hardest to control.
It is due to acoustics within the combustion chamber that often couples to the chemical combustion processes that are the primary drivers of the energy release, and can lead to unstable resonant "screeching" that commonly leads to catastrophic failure due to thinning of the insulating thermal boundary layer. Acoustic oscillations can be excited by thermal processes, such as the flow of hot air through a pipe or combustion in a chamber.
Specifically, standing acoustic waves inside a chamber can be intensified if combustion occurs more intensely in regions where the pressure of the acoustic wave is maximal. Screeching is often dealt with by detailed changes to injectors, or changes in the propellant chemistry, or vaporising the propellant before injection, or use of Helmholtz dampers within the combustion chambers to change the resonant modes of the chamber.
Testing for the possibility of screeching is sometimes done by exploding small explosive charges outside the combustion chamber with a tube set tangentially to the combustion chamber near the injectors to determine the engine's impulse response and then evaluating the time response of the chamber pressure- a fast recovery indicates a stable system.
For all but the very smallest sizes, rocket exhaust compared to other engines is generally very noisy. As the hypersonic exhaust mixes with the ambient air, shock waves are formed. The Space Shuttle generates over dB A of noise around its base. The sound intensity from the shock waves generated depends on the size of the rocket and on the exhaust velocity. Such shock waves seem to account for the characteristic crackling and popping sounds produced by large rocket engines when heard live.
These noise peaks typically overload microphones and audio electronics, and so are generally weakened or entirely absent in recorded or broadcast audio reproductions. For large rockets at close range, the acoustic effects could actually kill. More worryingly for space agencies, such sound levels can also damage the launch structure, or worse, be reflected back at the comparatively delicate rocket above.
This is why so much water is typically used at launches. The water spray changes the acoustic qualities of the air and reduces or deflects the sound energy away from the rocket. Generally speaking, noise is most intense when a rocket is close to the ground, since the noise from the engines radiates up away from the jet, as well as reflecting off the ground.
Then the largest portion of the energy is dissipated in the exhaust's interaction with the ambient air, producing noise. This noise can be reduced somewhat by flame trenches with roofs, by water injection around the jet and by deflecting the jet at an angle. Rocket engines are usually statically tested at a test facility before being put into production.
For high altitude engines, either a shorter nozzle must be used, or the rocket must be tested in a large vacuum chamber. Rocket vehicles have a reputation for unreliability and danger; especially catastrophic failures.
Contrary to this reputation, carefully designed rockets can be made arbitrarily reliable. However, one of the main non-military uses of rockets is for orbital launch.
In this application, the premium has typically been placed on minimum weight, and it is difficult to achieve high reliability and low weight simultaneously. In addition, if the number of flights launched is low, there is a very high chance of a design, operations or manufacturing error causing destruction of the vehicle.
The Rocketdyne H-1 engine, used in a cluster of eight in the first stage of the Saturn I and Saturn IB launch vehicles , had no catastrophic failures in engine-flights. The Pratt and Whitney RL10 engine, used in a cluster of six in the Saturn I second stage, had no catastrophic failures in 36 engine-flights. The Rocketdyne F-1 engine, used in a cluster of five in the first stage of the Saturn V , had no failures in 65 engine-flights.
The Rocketdyne J-2 engine, used in a cluster of five in the Saturn V second stage, and singly in the Saturn IB second stage and Saturn V third stage, had no catastrophic failures in 86 engine-flights. The Space Shuttle Solid Rocket Booster , used in pairs, caused one notable catastrophic failure in engine-flights. The Space Shuttle Main Engine , used in a cluster of three, flew in 46 refurbished engine units. These made a total of engine-flights with no catastrophic in-flight failures.
Rocket propellants require a high specific energy energy per unit mass , because ideally all the reaction energy appears as kinetic energy of the exhaust gases, and exhaust velocity is the single most important performance parameter of an engine, on which vehicle performance depends. Aside from inevitable losses and imperfections in the engine, incomplete combustion, etc. A triatomic molecule like water has six degrees of freedom, so the energy is divided equally among rotational and translational degrees of freedom.
For most chemical reactions the latter situation is the case. This issue is traditionally described in terms of the ratio, gamma, of the specific heat of the gas at constant volume to that at constant pressure. The rotational energy loss is largely recovered in practice if the expansion nozzle is large enough to allow the gases to expand and cool sufficiently, the function of the nozzle being to convert the random thermal motions of the molecules in the combustion chamber into the unidirectional translation that produces thrust.
As long as the exhaust gas remains in equilibrium as it expands, the initial rotational energy will be largely returned to translation in the nozzle. Although the specific reaction energy per unit mass of reactants is key, low mean molecular weight in the reaction products is also important in practice in determining exhaust velocity.
This is because the high gas temperatures in rocket engines pose serious problems for the engineering of survivable motors. Because temperature is proportional to the mean energy per molecule , a given amount of energy distributed among more molecules of lower mass permits a higher exhaust velocity at a given temperature.
This means low atomic mass elements are favoured. Liquid hydrogen LH2 and oxygen LOX, or LO2 , are the most effective propellants in terms of exhaust velocity that have been widely used to date, though a few exotic combinations involving boron or liquid ozone are potentially somewhat better in theory if various practical problems could be solved.
It is important to note in computing the specific reaction energy, that the entire mass of the propellants, including both fuel and oxidiser , must be included. The fact that air-breathing engines are typically able to obtain oxygen "for free" without having to carry it along, accounts for one factor of why air-breathing engines are very much more propellant-mass efficient, and one reason that rocket engines are far less suitable for most ordinary terrestrial applications.
Fuels for car or turbojet engines , use atmospheric oxygen and so have a much better effective energy output per unit mass of propellant that must be carried, but are similar per unit mass of fuel. Computer programs that predict the performance of propellants in rocket engines are available. With liquid and hybrid rockets, immediate ignition of the propellant s as they first enter the combustion chamber is essential. This is sometimes called a hard start or a rapid unscheduled disassembly RUD.
Ignition can be achieved by a number of different methods; a pyrotechnic charge can be used, a plasma torch can be used, [ citation needed ] or electric spark ignition  may be employed. Gaseous propellants generally will not cause hard starts , with rockets the total injector area is less than the throat thus the chamber pressure tends to ambient prior to ignition and high pressures cannot form even if the entire chamber is full of flammable gas at ignition.
Solid propellants are usually ignited with one-shot pyrotechnic devices. Once ignited, rocket chambers are self-sustaining and igniters are not needed. Indeed, chambers often spontaneously reignite if they are restarted after being shut down for a few seconds. However, when cooled, many rockets cannot be restarted without at least minor maintenance, such as replacement of the pyrotechnic igniter.
Rocket jets vary depending on the rocket engine, design altitude, altitude, thrust and other factors. Carbon rich exhausts from kerosene fuels are often orange in colour due to the black-body radiation of the unburnt particles, in addition to the blue Swan bands. Peroxide oxidizer-based rockets and hydrogen rocket jets contain largely steam and are nearly invisible to the naked eye but shine brightly in the ultraviolet and infrared.
Jets from solid rockets can be highly visible as the propellant frequently contains metals such as elemental aluminium which burns with an orange-white flame and adds energy to the combustion process.
Some exhausts, notably alcohol fuelled rockets, can show visible shock diamonds. These are due to cyclic variations in the jet pressure relative to ambient creating shock waves that form 'Mach disks'.
The shape of the jet varies by the design altitude: The Solar thermal rocket would make use of solar power to directly heat reaction mass , and therefore does not require an electrical generator as most other forms of solar-powered propulsion do. A solar thermal rocket only has to carry the means of capturing solar energy, such as concentrators and mirrors. The heated propellant is fed through a conventional rocket nozzle to produce thrust.
The engine thrust is directly related to the surface area of the solar collector and to the local intensity of the solar radiation and inversely proportional to the I sp. Ability to productively use waste gaseous hydrogen —an inevitable byproduct of long-term liquid hydrogen storage in the radiative heat environment of space—for both orbital stationkeeping and attitude control. Nuclear propulsion includes a wide variety of propulsion methods that use some form of nuclear reaction as their primary power source.
Various types of nuclear propulsion have been proposed, and some of them tested, for spacecraft applications:.
According to the writings of the Roman Aulus Gellius , the earliest known example of jet propulsion was in c.
The aeolipile described in the first century BC often known as Hero's engine essentially consists of a steam rocket on a bearing. It was created almost two millennia before the Industrial Revolution but the principles behind it were not well understood, and its full potential was not realised for a millennium.
The availability of black powder to propel projectiles was a precursor to the development of the first solid rocket. Ninth Century Chinese Taoist alchemists discovered black powder in a search for the elixir of life ; this accidental discovery led to fire arrows which were the first rocket engines to leave the ground. It is stated that "the reactive forces of incendiaries were probably not applied to the propulsion of projectiles prior to the 13th century".
A turning point in rocket technology emerged with a short manuscript entitled Liber Ignium ad Comburendos Hostes abbreviated as The Book of Fires. The manuscript is composed of recipes for creating incendiary weapons from the mid-eighth to the end of the thirteenth centuries—two of which are rockets. Heat Transmission. Melbourne, FL: Krieger Publishing, , pp.
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