Tallinn University of Technology. Introduction to Astronautics Sissejuhatus kosmonautikasse. Vladislav Pust õnski 2009. General characteristics. Liquid propellant rocket engines.
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Tallinn University of Technology
Introduction to AstronauticsSissejuhatus kosmonautikasse
The working principle of all liquid rocket engines is transformation of the potential chemical energy of liquid propellants to kinetic energy of the exhaust gases. It is important to mention that LRE is only a part of the propulsion system; other parts are tanks, plumbing, hydraulics, framework etc.
There are two basic types of LRE propellants: monopropellants and bipropellants.
Monopropellants are liquids which may be stored in a single tank (and remain stable). They are decomposed releasing energy in presence of a catalyst. Among such monopropellants are hydrogen peroxide H2O2 and hydrazine N2H4. Hydrogen peroxide is used, for example, as propellant for the turbopumps in RD-107/108 engines in the first and the second stages of the Soyuz launch vehicles. It was also used in the Mercury manned spacecraft, in the Centaur upper stage (in the ullage and attitude control motors) etc. However, this propellant slowly decomposes by itself, so it cannot be hold for years and thus cannot be used in spacecraft with long lifetimes. Hydrazine is widely used in maneuvering thrusters or main engines of spacecraft, also in descent engines. For example, hydrazine was used in the thrusters of the Voyager spacecraft, in the descent engines of the Viking and the Phoenix Martian landing probes etc. Hydrazine is stable and may be hold for years.
The main advantages of monopropellants is that they need only one tank and that the ignition system is not needed, they react by themselves in presence of catalyst. Due to low temperatures in the thrust chamber (combustion chamber) they may work for a long time (for hours) and be restarted (may give thousands of very short pulses). That makes them ideal for thrusters. Their thrust may vary from tens of grams to several kilograms. The main disadvantage is their low specific impulse (Isp), for hydrazine mostly Isp < 250 s.
Bipropellants consist of fuel and oxidizer that should be hold separately and react in the thrust chamber when ignited.
Bipropellant mixtures divide in practice by their properties into hypergolic and cryogenic propellants (although the first does not exclude the second). Hypergolic mixtures ignite spontaneously when brought in contact. Thus, no complex ignition system and starting procedure is needed (thus, multiple restarts are simply feasible). The process of combustion of hypergolic propellants is also more stable, so engines are more simple to develop and less likely to be destroyed at work. Hard starts are less likely with hypergolic propellants (hard start occurs when the ignition in the thrust chamber takes place at presence of excessive concentration of propellants, thus instantaneous overpressure is established and it may lead to explosive destruction of the engine).
The most widely used hypergolic mixture is hydrazine or its variants (monomethylhydrazine /MMH/,unsymmetrical dimethilhydrazine /UDMH/, aerozine /50% hydrazine + 50% UDMH/) as fuel and dinitrogen tetroxide (N2O4) (or nitric axid /NO/ in earlier applications) as oxidizer. The great advantage of this propellant is that the fuel and the oxidizer both are liquids at the normal conditions, they are not cryogenic, so they may be stored at the common temperatures. This makes them ideal for military applications (ICBM may stay for arbitrary time in its silo or stored) as well as for spacecraft with a long lifetime (like interplanetary probes). This is the reason why they have been used on many ICBMs which later have been converted into launch vehicles: the Titan, the Strela and the Rokot (ICBM UR-100/100H), the Dnepr (R-36M). These mixtures have quite high Isp (vacuum values 300 – 320 s and even more), so they are most common on spacecraft, although need more complex engines when that based on the hydrazine as monopropellant. The Cassini spacecraft uses hydrazine monopropellant thrusters for small attitude maneuvers, but burns
N2O4/MMH in its main engine. Orbital engines of the Space Shuttle use N2O4/MMH, Apollo LM and CSM used N2O4/aerozine, Luna E8 series (Lunokhod, soil sample missions) used N2O4/UDMH. The highest disadvantage of this mixture is that it is highly toxic (both fuel and oxidizer), so it should be handled very carefully. The flame of hypergol is nearly colorless, slightly blue. Plumes of even powerful engines burning this propellant is nearly invisible.
Cryogenic propellants include mixtures at least one component of which (fuel or oxidizer or both) need low temperatures to liquefy. One of the most common such mixtures is LOX/kerosene (some special technological processes are used to produce kerosene for rocket industry, the resulting fuels have different names; in US RP-1 is a popular type of kerosene fuels, in USSR sintin was used). LOX/UDMH is also an option. LOX, one of the most widespread oxidizers, boils at –1830C, thus being moderately cryogenic: this temperature is higher than the boiling point of air (– 1940C), which is produced in industrial quantities. So, air does not liquefy on cold walls of LOX tanks, and extensive termostating is not needed. For LOX/kerosene propellants, in vacuum Isp 340 s, and it is very widespread propellant for launch vehicles, specially for the first stages. It is used on all stages of the Soyuz and the Zenit, on the first stage of the Atlas V, it was used on the first stages of the Saturn I/IB, the Saturn V and the Energia. Due to its better energetic characteristics than that of the hypergols, LOX/kerosene is also used on the Blok DM, the 4th stage of the Proton (3-stage N2O4/UDMH heavy rocket). Cryogenic propellants cannot be held for a long time, and the tanks should be insulated. In space, some amount of the propellant boils out, and if some time passes between multiple burns, the loss may be significant, so instulation is needed. Plumes of LOX/kerosene engines contain many carbon particles and thus are bright yellowish.
The most energetic propellant that is used in practice is LOX/LH2, it may have Isp > 450 s (vacuum value). Liquid hydrogen is highly cryogenic fuel, having the boiling point at –2530C.
So it is difficult to use, since continuous thermostating is unavoidable. Tanks containing LH2 should have extensive thermal protection (which leads to increase of their weight), the fuel cannot be hold in space for a long time since it boils out. Extremely low temperatures of LH2 lead to changes of physical properties of metals which contact with this fuel, metals saturate with H2; these factors should be taken into account at building of a LOX/LH2 rocket engine, plumbing and tanks. Because of very low density of LOX/LH2 (~280 kg/m3), it needs large tanks, which also increase the mass of stages with this propellant. In addition, Isp of such engines significantly drops in the atmosphere. So, LOX/LH2 is used primarily on the upper stages of the launch vehicles. The first stage where this propellant was applied was the Centaur upper stage for the Atlas launch vehicle. Later it appeared on the upper stage S-IV (Saturn I) and S-IVB (Saturn IB and Saturn V), as well as on the second stage S-II of the Saturn V. The Space Shuttle has become the first spacecraft where LOX/LH2 is used on the stage working from the sea level. Later it appeared on the second stage (also working from the sea level) of the Energia. Today it is a common propellant on upper stages of launch vehicles: Delta, Atlas, Ariane, GSLV (India), CZ-3 (China). It is also used on the first stage started on the sea level, but is aided by strap-on solid rocket boosters: Ariane, H-II (Japan). There is only one full-cryogenic launch vehicle which burns this propellant on all stages, that is the Delta IV. The plumes of LOX/LH2 engines are nearly invisible.
To introduce propellants into the thrust chamber, two principle designs of LRE are used: pressure-fed and pump-fed. The first one is the most simple and reliable, the second one enables to get higher specific impulse.
In a pressure-fed LRE, the propellants are forced to the thrust chamber by pressure of gas which pressurizes the tanks. For pressurization, a separate gas supply is provided. So, there is a special tank with pressurizing gas onboard (helium is commonly used for this purpose).
The greatest advantage of pressure-fed engines is simplicity and thus reliability of this design: contrary to pump-fed engines, no complex turbopumps are needed, no gas generators etc. Such engines contain much less parts and much less moving parts, so there are much less things that might fail. The procedure of engine cut-off and restart is also very simple: there is no need to stop and restart the turbopumps, its enough to close or open the valves, and the propellant flow to the thrust chamber ceases or recommence (if the propellant is hypergolic, even no ignitor is needed). To avoid the pressurizing gas to cool down due to expansion inside the tanks, it is often warmed up in the heat exchanger.
The advantages of this solution make pressure-fed engines ideal for applications where reliability and simplicity are important, as well as capability for multiple restarts. This is the reason why all engines of the Apollo CSM, as well as all engines of the Apollo LM were pressure-fed. Shuttle orbital maneuvering and control engines are pressure-fed as well. Maneuvering and attitude control thrusters of satellites and space probes are mostly pressure-fed since they are restarted thousands of times.
However, this design have two principle disadvantages (mutually related). Specific characteristics of rocket engines depend on the pressure in the thrust chamber. But the
pressure in the thrust chamber cannot exceed the pressure in the the tanks (actually, the pressure in the tanks should be higher). To withstang high pressure, large tanks should have more robust and heavier construction. The larger is the tank, thicker should be its walls to bear the same pressure. Thus, pressure-fed systems are generally limited by chamber pressures of ~10 bar (on the Apollo SM it was 7 bar, on the LM Ascent Stage 8.4 bar). They are rarely applied on first stages due to large size of the tanks (however, the engine on the 2nd stage of the Delta II rocket is pressure-fed). To avoid additional pressure in the tanks, regenerative cooling jacket is often avoided, that obliges to use ablative and radiative cooling.
Pump-fed systems do not have the limitations of the pressure-fed systems. In this design, the propellant is forced into the thrust chambers with dedicated pumps. The required efficiency may be provided only with centrifugal pumps, herewith the pump should rotate at tens of thousands rpm. Only a turbine is capable to ensure such speeds, so the natural solution is a turbopump. A turbopump consists of one or more pumps often mounted on the same shaft with a driving turbine. The turbine is driven by gas flow, the gas may be produced in a gas generator by preburning some amount of the propellant, by burning a separate propellant (like hydrogen peroxide in the RD-107/108 engines on the Soyuz) or by gasification of some propellant in the cooling jacket of the thrust chamber and the nozzle. The pumps may be multistage. The turbopump assembly may include also booster pumps, which are added to unload principle pumps and to increase the pressure in the gas chamber (these pumps may be driven by a hydraulic turbine powered by liquid from a high-pressure line, but also by the main turbine). Turbopump assembly is the most complex part of the engine, since it should
have enormous productivity and work in harsh conditions (the turbine is driven by very hot gases and rotates very quickly). For example, the turbopump of RD-170/171 (the most powerful LRE ever produced, Energia &Zenit launch vehicles, LOX/kerosene) has a mass flow rate of ~2.4 tons/s, it provides the pressure in the thrust chamber of ~250 bar, the power of the turbine is ~200 MW, it rotates at ~14 000 rpm. The pressure of gases driving the turbine is ~500 bar, their temperature is ~5000C. At the same time the turbopump should be compact and lightweight (the mass of whole RD-170 is about 10 tons). So high characteristics are possible only because the lifetime of such assemblies is only tens or hundreds of seconds. However, there exist pump-fed engines of multiple use which may work for hours, be restarted and continue to work after revision. An example of such engines is the Space Shuttle Main Engine (SSME).
The obvious advantage of pump-fed engines is that they may provide very high pressures inside the thrust chamber and so their specific impulse is high. In spite of their complexity, they may be compact enough and be lighter than pressure-fed engines with their pressurizing gas vessels and thick propellant tanks; thanks to their efficiency, they make it possible to spend less propellant. Their complexity is the highest disadvantage, since complex turbopump assemblies tend to be more expensive and less reliable than pressure-fed designs. However, if efficiency is critical, pump-fed design is a natural solution. Turbopumps are used on all stages of launch vehicles, but also on spacecraft. The Soviet lunar probes E8 (Lunokhods, soil sample missions) used pump-fed design, and the Soviet lunar module for the manned expeditions as well (since weight was critical). The space stations Salyut, the Soyuz manned spacecraft have been provided with pump-fed engines.
There are several designs of pump-fed engines. The most spread is the gas generator cycle. In these engines the turbine of the turbopump is powered by gas resulting from burning some of propellant in the gas generator (also called preburner sometimes) – a special small combustion chamber. In some designs there may be two gas generators (like the RD-170/171), sometimes each gas generator provides gas for separate turbines (of fuel and oxidizer). The mixture in the gas generator is ordinarily very fuel-rich or oxidizer-rich in order to keep the temperature reasonably low and not to damage the blades of the turbine (actually, only small amount of the propellant burns, the rest is only gasified). After the turbine, the gas is ejected, either through the main nozzle either through a special nozzle. Due to its low temperature, its contribution to the engine thrust is quite low, so it is nearly “wasted” for the thrust. Several percent of the propellant are lost. However, sometimes this gas is used in steering nozzles or may participate in film cooling of the main nozzle (like in the F-1 engine of the Saturn V).
To improve efficiency of the engine, another version of this cycle is used, that is the so-called staged combustion cycle (or closed cycle). The main difference of this cycle is that the gas after turbine is not dumped, but is returned to the thrust chamber. So, all propellant and all heat pass through the thrust chamber and nothing is wasted. The disadvantage of this solution is that the turbine have to do work against the pressure of the gases which it should press into the thrust chamber. So the efficiency of the turbine drops, and it needs more power to work. Thus, it works in worse and more harsh conditions, the plumbing of hot gases ducts is much more complex, as well as the control. So such engines are generally more complex, more expensive and less reliable. They are very sensitive to productional quality and to external particles that may occasionally get into the ducts, turbines and pumps. But the gain of Isp may be so high that this design makes sense. It first appeared in the USSR, and they have a long tradition of building engines of the closed cycle. For example, the RD-170/171 applies the
closed cycle (contrary to the F-1), as well as the SSME of the Shuttle.
In most cases only small amount of the propellant is gasified in the gas generator (and in the gas generator cycle it cannot be else to avoid excessive loss of propellant). But in some developments the full amount of the fuel and the oxidizer passes through the turbine (the so-called full flow staged combustion cycle). It enables to reduce the temperature of the gas and the rotation velocity of the turbine, since the it is driven by larger mass. The lifetime and reliability grow. Of course, two separate gas generators and turbines are needed for the fuel and the oxidizer. However, separate systems for both components are usual for LOX/LH2 engines, since the components have very different physical properties (density on the first place), so it is difficult to provide optimal characteristics for them in a single assembly.
Sometimes it is possible to get rid of the gas generator assembly at all (gas generator is a small combustion chamber by itself, with its own nozzle ejecting gas into the turbine, so it is a complex unit). This is the expander cycle design. In this cycle the gas for driving the turbine is produced from the fuel vaporized in the cooling jacket of the thrust chamber and the nozzle. A gas generator is sometimes used to start the engine. This cycle may be opened or closed. In the opened cycle, only a small portion of the fuel is used to drive the turbine and thereafter it is dumped. In the closed cycle the fuel is redirected into the thrust chamber after leaving the turbine. Although the close cycle saves fuel, the open cycle enables higher pressure drop on the turbine which increases its efficiency and enables to raise the pressure in the chamber. This leads to higher Isp (this is the case of LE-5A/B on the second stage of the Japanese H-II rocket, LE-5 used the gas generator cycle). The famous RL-10 and its modifications on the Centaur upper stage use the expander cycle. Generally, the expander cycle is mostly applied in LOX/LH2 engine since fuel is ordinarily used for regenerative cooling (oxidizer is too reactive) and LH2 has low boiling point and is very effective as reaction mass.
The thrust chamber is the principle component of the rocket engine, the propellant is injected into it and burns, transforming into hot gases that escape through the nozzle (the bell). The thrust chamber assembly consists of the following main components: the thrust chamber body, the nozzle (the chamber narrows to its end, the narrowest part of it is called throat, and behind the throat it expands again, forming the nozzle; the end part of the nozzle is called extension), the injector. In the case of regenerative cooling, the body and the nozzle may be combined with a cooling jacket.
Cooling of the thrust chamber and the nozzle
Since gases in the thrust chamber have very high temperatures, its walls should be cooled, as well as the walls of the nozzle. Without cooling, the walls cannot withstand such temperatures for a long time. There are two principles of cooling: passive and active cooling. Different methods may be applied simultaneously in different parts of the chamber and the nozzle.
Passivecooling includes ablative and radiative methods. Ablative cooling means that the walls are covered with substance called ablation, which have high heat capacity and absorbs heat by transforming itself chemically and/or physically. The ablation burns slowly and removes heat with the gases created in this process. This method is limited by timespan and by temperature: ablative materials cannot withstand very high temperatures and since they are gradually removed, ablation works only for a limited time. However, the highest advantage of this method is simplicity and reliability, so it is sometimes applied even on large engines, like RS-68, the most powerful LOX/LH2 engine in use (Delta IV).
Radiative cooling is the process when the hot wall loses heat by radiation. Being very
simple, this method is limited to thin surfaces with relatively moderate incident heat fluxes. If the heat flux is very intensive, the equilibrium temperature of such wall with only radiative cooling is too high and the wall may be damaged. If the wall is thick, its hot side is damaged before the heat diffuses to the cold side. So, radiative cooling is mostly applied to cool nozzle extensions, as well as in small maneuvering engines, where the heat of short burns is absorbed by a massive conductive wall of the chamber (made from cooper alloy, for instance) and is irradiated between the burns.
Active cooling methods include regenerative cooling and film cooling. Regenerative cooling means that a flow of cold propellant is organized along the the hot wall and the propellant carries away excessive heat. Thus, the walls of the thrust chamber and the nozzle have a cooling jacket with a propellant flow inside (fuel is commonly used). The propellant is directed into the jacket from the tank before it is injected into the thrust chamber. There are different ways to build the cooling jacket. The simplest way is two walls, inner and outer, separated by a folded metal sheet, with propellant flowing along the folds. This design have been preferred in Russia and now is also applied in US. The liquid may also flow along rectangular channels machined or formed into a liner fabricated from high-conductivity material (like cooper alloys). The example of such design is the SSME. In US the traditional design have been the thrust chamber and nozzle built from thin rectangular tubes strengthened by outer bracing (tubular wall). The tubes are directed downwards and upwards the walls of the thrust chamber and the nozzle. Since the diameter of the nozzle changes along its axis, the form of the tubes also change (but their cross-section remains constant). The tubes may bifurcate. The hot wall is very thin, so heat exchange with the liquid is very effective. The example of this design is the F-1.
Efficiency of the regenerative cooling is very high, but there are also some limitations. In
the throat, the diameter and thus the surface of the chamber wall is small, and it is impossible to pump enough liquid along the limited surface to provide cooling. It is not always possible to increase the velocity of the flux since it would require raise of pressure that forces the propellant through the cooling jacket. In this case, film cooling is additionally applied. The main idea of film cooling is that some fuel in injected through additional injectors into the hotest parts of the thrust chamber right against the wall. The liquid fuel absorbs heat by boiling and evaporating, thus a protective cold boundary film is created and protects the wall from contact with the hot gas. This film is spread along the wall by gas moving along the chamber. A variation of this method is transpiration: the coolant gets into the chamber from the jacket through a porous chamber wall. Film cooling may be realized also with cold gas from turbine directed along the wall of the nozzle to protect it from hot gases from the thrust chamber. The protection of the nozzle extension of the F-1 engine was performed in this way.
The propellant is introduced inside the thrust chamber through the injector. The injector forms a spray of the components to provide their effective mixing an burning. A typical injector head consists of a plate with holes for the propellant components organized in a special pattern. Some of the injector elements may represent sleeves sticking out from the plate, they may have multiple holes. The injector head may also be divided into sections by partitions.
Stability of the combustion process in the chamber depends highly on the effectiveness of mixing and thus, on the injector. The size of propellant drops and the parameters of the spray define the lifetime of the drops, intensity of their evaporation and the quality of the burning mixture. There is a number of reasons why burning process may become unstable, and the combustion may become resonant. For example, pressure pulsation may influence the
injection system: raise of pressure inside the chamber slows down the injection rate, and the following drop of pressure (when the excess of the propellant leaves the chamber) leads to a new increase of the injection rate. Self-oscillating process thus establishes with the frequency from tens to hundreds cycles per second. This instability is in strong dependence on the lifetime of propellant drops, i.e. on the delay between the propellant injection and its combustion. This instability is often eliminated by changing pressure drop on the injector; an injector pressure drop usually makes ~1/4 of the chamber pressure.
Another type of instabilities is high-frequency combustion instability (frequencies > 500 cycles per second), it is the most dangerous and is specially pronounced in engines having high thrust. These instabilities arise because the time of drops vaporization is not constant and depends on the pressure near the injector head. The higher is this pressure, more intensively vaporize the drops, the combustion process accelerates, and a shockwave spreads along the chamber, reflects from the opposite wall and returns, raising the pressure even more. The period of the oscillations depends on the time in which the shock front returns. These oscillations usually destroy a thin-wall chamber within seconds. This kind of oscillations is suppressed by changing the chamber length and width, by installing additional partitions inside the chamber which divide it into smaller volumes, and by matching the injector head (number and position of the holes, the sleeves etc.) The designers of the F-1 engine, the biggest one-chamber engine ever used, faced the high-frequency instabilities and had many problems with them. The problems were solved by matching the injector head: it was divided into sectors by partitions. Vaporization of the components before introducing them into the chamber also may be applied. A radical way to solve the problem is to replace one large thrust chamber by several smaller ones: due to smaller dimensions of the single chambers, the danger of appearance of high-frequency instabilities is smaller. This way was chosen by the designers of RD-170/171: the engine has 4 chambers with the thrust of ~200 tons per
chamber. Smaller size of the chamber also permits to decrease the length of the engine. High-frequency instabilities make creation of high-thrust LREs quite problematic and expensive.
There is another instability, low-frequency combustion instability (typically 10 – 100 cycles per second), not directly related to the injector. It appears due to resonance between the thrust caused by changes of the fuel flow rate, and and proper frequencies of the tanks and structure of the rocket. The result is so-called pogo oscillations of the rocket. This kind of oscillations is often eliminated by dampers in propellant lines, like in the Space Shuttle and Saturn V. To damp oscillations, a small amount of helium is introduced into the propellant line to shift the natural frequency of the line and to destroy the resonance.
In general, LREs are throttled by adjustment of the amount of the propellant delivered into the thrust chamber, and this idea may seem to be quite simple. However, in practice it is not so simple to realize. If the engine is pump-fed, we should take into account the fact that the turbopumb is driven by the propellant, and if the flow rate is decreased, there may be lack of reaction mass to power the turbine. The pumps are also designed for a certain propellant flow, and significant change of the flow rate may lead to significant drop of efficiency, and the turbopump assembly would work in unbalanced conditions: the power of the turbine may be insufficient to drive the pumps. Of course, these problems may be overcome technically, but that would mean impossibility to provide optimum conditions for the turbopump: as everywhere in technical sciences, an universal unit is usually not so effective through the whole wide range of working conditions as an specialized unit could be.
When the problems of the turbopump are solved (or in the case of a pressure-fed engine), problems with cooling jacket may arise (if cooling is regenerative). Drop of propellant flow rate may slow down and liquid may begin to vaporize, thus the engine would stall. The amount of the coolant available also would drop. The temperature in the thrust chamber will remain nearly unchanged if the mixture ratio is intact, and it may become impossible to cool down the walls. A possible solution is to change mixture ratio and thus to decrease the temperature in the thrust chamber, but this would lead to decrease of Isp (which would drop anywhere since the pressure in the chamber will drop). Another solution is to use only passive cooling methods (ablation), but that would mean lower temperatures and a shorter lifetime of the engine.
But the turbopump and cooling are not the only issues to take into account. The mayor problem is combustion stability. When the propellant flow rate decreases, the injector pressure drop falls quicker than the pressure inside the chamber, and, as we have seen, the flow rate becomes dependent on small variations of pressure in the chamber. A feedback between the chamber pressure and the propellant supply appears and combustion becomes unstable. To avoid that problem, variable-geometry injectors may be used: the area of the injector head is decreased when the flow rate drops (this was the case of the Apollo LM Descent Stage).
Generally, most of LREs may work with moderate throttling by several percent without stalling nor serious loss of efficiency. But deep throttling requires special designs and is a difficult problem to solve. The highest throttling range among the human-rated engines was the TRWS for the Descent Stage of the Apollo LM. It could be throttled down to 10%. However, the range of ~65% 95% was unusable due to stability issues. SSME is throttled in-flight in the range of 65% 105% (and a little bit wider range is available). RD-180 (a two-chamber version of RD-170) is throttled in the range of ~40% 100%.
To start a LRE, several operations should be performed in a right sequence. First of all, the tanks should be pressurized. This is done with separate gas (like helium stored in special vessels) or vaporized propellant components. In the last case the vaporized components are available only after the engine is started, so temporarily some other gas is used. If cryogenic components are used, specially LH2, the plumbing should be chilled with a small initial flow of the component before the full flow is opened. This is done to prevent boiling of the cryogenic component inside the plumbing.
If turbopump is applied, the turbines should be started to begin delivery of the propellant into the chamber. This may be performed by burning main components in the gas generator: the initial amount of propellant is directed into the gas generator to gasify. Turbines may also be started by initial flow of a separate gas like helium stored in a starting vessels. When turbines are able to pump components, the flow of the propellant into the chamber is initiated by opening main vents.
If propellants are not hypergolic, they should be ignited. For a single-burn engines, like that of launch vehicles, chemical ignitors are often used : a small amount of hypergolic propellant is introduced into the gas generator and the thrust chamber before the main propellants, and they ignite the propellant mixture (often this hypergolic component self-ignites when mixed with the main oxidizer). Ignition with a pyrotechnic charge or even a torch introduced into the thrust chamber also may be applied. For multiple-burn engines electric ignition (with spark) is often used.
To avoid hard start (when the pressure in the thrust chamber rises too quickly, damaging the chamber), the start of a large engine should be performed carefully. Sometimes the start is
performed in two stages: at first the engine is started at a fraction of the full thrust, and when the thrust is raised to the nominal value. The engine may also be started at a mixture ratio different from the nominal, and then the ratio is set to the nominal.
Large engines cannot be cut-off by only closing main propellant valves. Since turbines continue rotating for some time, the pressure in the pump tract will not drop immediately. To avoid raise of pressure at the main valves, the flow may be redirected to a low pressure line. If the engine is pump-feed, at first the propellant flow to the gas generator is decreased, and the turbines rotation slows down. Large engines are often cut-off in two stages to avoid rapid transient processes that might damage the engine and the rocket. Pressure-fed engines may be cut-off by propellant depletion: the flow of one of the components stops, and the engine shuts down itself.
In the previous lecture we have seen that the thrust T of a rocket engine is defined by the formula
Here q = dm/dt – mass flow rate, S – area of the nozzle, pa – static pressure of the exhaust gases, p – ambient air pressure. If pa > p, the nozzle works with underexpansion, if pa < p, the nozzle works with overexpansion, pa = p is the case of ideal expansion. Increase of the expansion ratio (the ratio of the pressure inside the chamber pch and the pressure at the nozzle exit pa, =pch/pa) leads to an increase of the first summand and to a decrease of the second summand. It may be proved that the case of ideal expansion provides the highest thrust possible for the corresponding external pressure. As the rocket ascends, the ambient pressure changes, so it is impossible to provide optimal conditions for all heights. A nozzle that has overexpansion at lower heights (after the lift-off) will work in underexpansion conditions after the liftoff. In vacuum, every nozzle works in underexpansion conditions, since it is impossible to provide zero pressure at the nozzle exit (for that, the nozzle expansion should have infinite length and with). But larger is the nozzle, less is the underexpansion. However, it makes no sense to increase very much the size of the nozzle, since the gain of efficiency would be cancelled by grow of the nozzle size and weight. For the first stage, the expansion ratio of the nozzle cannot be very large, since too high overexpansion would lead to separation of the jet of the exhaust gases from the nozzle. That may lead to catastrophic drop of efficiency and even damaging the nozzle, if measures are not taken. Generally, the nozzle of the first stages work in overexpansion conditions in the lower atmosphere. Nozzles of upper stages work at significantly lower ambient pressures, so they have much larger expansion ratios than the
nozzles of the first stages. The expansion ratio of a nozzle is a compromise between thrust efficiency at different heights, weight and size. In the recent time an original design have been implemented in several engines: the nozzle extension changes inflight to improve efficiency while the rocket ascends from the denser atmosphere to more rarefied layers. This solution is applied on the second stage of the Delta IV rocket (LOX/LH2 engine RL-10B-2).
The shape of the rocket nozzle is of de Laval type. It is convergent-divergent, having a throat. In the convergent section of the nozzle the gas flow is subsonic, it velocity increases towards the throttle and achieves the sound speed at the throttle. In the divergent section the gas jet expands and achieves supersonic speeds.
The gas escapes in slightly different directions in different parts of the nozzle. This fact
may be taken into account with the coefficient that relates the Vmeanescape may velocity to the actual escape velocity Va and the half-angle of conicity of the nozzle :
From gas dynamics, the gas escape velocity may be found as
R – gas constant, –expansion ratio, Tch – temperature inside the
thrust chamber, – adiabatic index, – molar weight of the exhaust gases, k – imperfection coefficient that takes into account incomplete combustion, film cooling etc.
It is seen that higher escape velocity are provided with higher temperatures in the chamber and lower molar weights of the exhaust. The dependence of the expansion ratio has a
maximum at = 0, i.e. the pressure inside the thrust chamber should be possibly high and the outer pressure should be possibly low.
The term in the brackets is – expansion efficiency, Vc = (RTch / )1/2is the sound speed. So formula the escape velocity may be given a form
For typical propellants = 1.25, and the escape velocity is ~2.8 Vc (if k~ 1 and ~ 1).
Launches of Proton & Titan II
Launch of Proton. 6 engines, N4H4/UDMH. The plume is nearly invisible (By source)
Launch of Titan. 2 engines N4H4/aerozine. (By source)
Launches of Soyuz & Saturn IB
Launch of a Soyuz. 5 engines, LOX/kerosene. The plume is yellowish bright(By source)
Saturn IB launch. 8 engines, LOX/RP-1 (By source)
Launches of Ariane V & Delta IV
Launch of Ariane V. 2 SRB + 1 LOX/LH2. The plume of the LOX/LH2 is nearly invisible(By source)
Delta IV Medium launch. 1 engine, LOX/LH2. The color is caused by nozzle ablative covering (By source)
Scheme of the pressure-fed engine
MA-5 turbopump assembly
Turbopump of MA-5 engine for the Atlas.(By source)
Scheme of the gas generator cycle
Scheme of the staged combustion cycle
Scheme of the expander cycle
1) injector; 2) thrust chamber body; 3) nozzle part; 4) throat; 5) holes; A, B) propellant components.(By V.I.Feodosjev, Osnovy tehniki raketnogo poleta.)
1) outer jacket; 2) inner cooled jacket; 3) soldering; 4) cooling liquid flow; 5) hot gases; (a) typical size several mm.(By V.I.Feodosjev, Osnovy tehniki raketnogo poleta.)
Outer wall of F-1 engine for the Saturn V.(By source)
Cooling jacket of F-1
Inside view of the F-1 engine (Saturn V). Bifurcating tubes are clearely visible(By source)
Injector head of the HM-7 engine (LOX/LH2, Ariane upper stages). Sleeves for one of the components are inside the holes for the another component(By source)
Injector head of the RD-170/171 engine (LOX/kerosene, Energia/Zenit first stage). Sleeves for one of the components form partitions(By source)
F-1 and RD-170
F-1 (vacuum thrust ~800 tons)(By source)
4-nozzle 4-chamber RD-170 (vacuum thrust ~810 tons)(By source)
Nozzle expansion at different heights
Diagram of the velocity V, pressure P and temperature T in a de Laval nozzle. M is the Mach number
de Laval nozzle