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[1.0] Chemical Rocket Systems

v1.4.1 / chapter 1 of 7 / 01 may 22 / greg goebel

* The primary rocket engine technology for spaceflight propulsion at the present time is the chemical rocket engine, and it is likely to remain the dominant technology for some time to come. This chapter provides an introduction to chemical rocket engine systems.

Thor LR-79 / MB-3 engine


[1.1] BASIC CONCEPTS OF ROCKET PROPULSION
[1.2] LIQUID PROPELLANT ENGINES: FUNDAMENTALS
[1.3] LIQUID PROPELLANT ENGINES: A SURVEY / AEROSPIKE & RBCC ENGINES
[1.4] SOLID PROPELLANT ENGINES: FUNDAMENTALS
[1.5] SOLID PROPELLANT ENGINES: A SURVEY / NEXT-GENERATION SOLIDS

[1.1] BASIC CONCEPTS OF ROCKET PROPULSION

* All rocket vehicles work on the principle of reaction, or "recoil", which is a consequence of the law of conservation of momentum. If a cannon fires a cannonball, the cannonball flies away with a momentum equal to the mass of the cannonball times its velocity. The shot gives the cannon the same momentum in the opposite direction, and if it were free to move without interference from friction or other constraint, it would fly backward, with a velocity less than that of the cannonball by the same factor that the cannon's mass is greater than the cannonball. For example, if the cannon weighs 25 times more than the cannonball, its velocity will be 25 times less than that of the cannonball.

There's a popular misconception that a rocket vehicle "pushes against the air", but that's not the case. It pushes against itself, using the recoil of the hot gases shot out the rocket engine exhaust to drive itself forward, and the air just gets in the way, causing drag and friction. The heavier the mass of the exhaust flow and the greater the velocity of that flow, the greater the recoil generated by the rocket engine, and the greater the thrust.

There are various ways to generate this thrust, though in all cases the result is the same, to expel a gas at high velocity. Nuclear rocket engines run a fluid through a nuclear reactor. Electric rocket engines typically accelerate ions to high velocities using electrified grids. Chemical rocket engines, the subject of this chapter, burn a "fuel" and an "oxidizer", either in a solid mixture or stored as liquids in separate tanks, and blast the exhaust out a usually bell-shaped ("convergent-divergent" or "con-di") nozzle.

Rocket engine thrust is formally measured in newtons (N) in the metric system; in pounds force (lbf) in the English system; and sometimes in kilograms force (kgp, where the "p" stands for the French "puissance / force"). Since the kilogram is a measure of mass, not force, kgp is a little dubious from the strict physics point of view, but it is equivalent to newtons divided by 9.81, and at least at one time was a fairly common measure of thrust.

Efficiency of a rocket engine can be measured in terms of exhaust velocity, but since the actual thrust is also dependent on the mass of the exhaust gas, a more useful measure is "specific impulse (Isp)", or thrust produced by a unit mass of propellant per second. In metric units, Isp is defined as "newtons per kilogram of propellant per second", and in English units it is defined as "pounds thrust per pound of propellant per second". The second definition, by the way, evaluates to "seconds", and that's normally how specific impulse is described. Specific impulse can be thought of as an index of the "mass ratio" of a rocket vehicle, or the ratio of payload to vehicle mass: the higher the specific impulse, the greater the efficiency in terms of the amount of payload per fuel mass.

However, this is a somewhat narrow definition of "efficiency", a much more practical one being how much payload can be lifted at a given cost, and in such terms an engine with the highest specific impulse may not be the most efficient. In addition, high specific impulse does not necessarily mean high thrust, and in fact as later chapters will show, highly efficient engines with high values of specific impulse tend to have very low thrust.

* A rocket vehicle obviously consists of some sort of airframe or casing mounting a rocket engine and providing storage for propellants. Less obviously, it must also carry guidance and control systems.

Holiday firework rockets simply have a stick or fins to keep them flying straight. Unguided rocket projectiles used by combat aircraft to attack ground targets or launched as a form of "barrage" artillery also generally use fins, though they also may have multiple rocket exhausts canted at an angle around the centerline of the rocket to cause them to spin for stabilization.

For large rocket vehicles, such as long-range missiles or space launch boosters, which are the focus of this discussion, more sophisticated control schemes are required. Some large rocket vehicles do have fins -- the Soviets developed a slick "grid fin" or "lattice fin", like a honeycomb paddle that pops out from the body of the rocket, with the trick catching on elsewhere. However, fins are only useful at low altitudes, since they are ineffective once the vehicle leaves the atmosphere. There are several approaches for control of such large vehicles:

Most modern large rocket vehicles are actually assemblages of separate rocket vehicles, known as "stages", that are stacked on top of each other. A big rocket vehicle contains a large amount of fuel, and as the fuel is burned up, the vehicle carries more and more useless dead weight. With "staging", when one stage is exhausted, the next ignites and the first is discarded.

Staging is a tricky operation, in essence trying to launch one rocket vehicle on top of another while the whole assembly is in flight, and some early long-range missiles used dodges to simplify the scheme. The original American Atlas missile, for example, used "half staging", with a three-engine assembly in a skirt at the bottom; two of the engines were discarded along with the skirt after initial boost, leaving a single sustainer engine.

The Soviet R-7 missile / SL-1 booster used "clustering", with four auxiliary boosters clustered around a similar central "core" booster, the auxiliary boosters being discarded after initial boost. Both these schemes allowed all the engines to be ignited at take-off, simplifying launch procedures. They were perfectly practical rocket vehicles, and in fact their descendants are still in use.

More modern large rocket vehicles use true staging, though they often also use a set of auxiliary solid-fuel or liquid-fuel "strap-on" boosters attached to the first stage to provide additional thrust. Each stage is connected to the others by a "shroud", or in the case of many Soviet-Russian missiles and boosters, by an open framework. The open framework has the advantage that the upper stage can be ignited before the lower stage is discarded, which simplifies the staging process. The Soviets also came up with an ingenious staging scheme for their R-27 / SS-N-6 submarine-launched ballistic missile to produce a more compact vehicle, with the engines for the payload stage actually contained in the fuel tank for the main stage. Pyrotechnics were used to cut open the tank so the rocket engine could ignite.

* While the lower stages of a multistage rocket vehicle are discarded before the machine leaves the atmosphere, the upper stages and the spacecraft they carry, if there is a distinction, need to operate in space. Such "space vehicles" generally need "restartable" rocket engines that can be turned on or off, which is a somewhat tricky problem because under "zero-gee" conditions the propellants do not tend to flow to the bottom of the tanks. To change its orientation, a space vehicle also needs a "thruster" systems, which are small rockets fitted into a vehicle to control its roll, pitch (nose up and down), and yaw (nose side to side) orientation, as well as precision acceleration and braking for tasks such as space rendezvous.

In addition, a practical rocket vehicle must contain a number of support systems, such as a radio "command and telemetry" system to allow ground controllers to send commands to the vehicle and receive operating status; sensors and diagnostic systems to indicate the vehicle's health; and a "self-destruct" system to blow the vehicle out of the sky should it suddenly go off course and threaten to fall on a populated area. However, such details are beyond the scope of a document on space propulsion systems.

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[1.2] LIQUID PROPELLANT ENGINES: FUNDAMENTALS

* Rocket researchers investigating spaceflight in the 1920s and 1930s knew that old-fashioned fireworks rockets weren't the way of the future for their work, and turned instead to liquid propellants to build the large rocket vehicles they wanted. By the mid-1960s, they had developed liquid-propellant rocket engines powerful enough to put men on the Moon.

In its simplest form, a modern liquid propellant rocket engine system consists of a tank of oxidizer and a tank of fuel that feed a combustion chamber, with the combustion exhaust forced out a bell-shaped nozzle. Three classes of liquid propellants are in common use:

Of course, a rocket is a very sophisticated piece of hardware, and the simple description above glosses over many fine details, some of which are far from obvious.

The propellant tanks are usually stacked on top of each other. In such a stacked configuration, the output line for the top tank is routed through the bottom tank. The outlets of the propellant tanks are closed by valves until the rocket engine is fired.

The simplest way to get the propellants into the combustion chamber is to use a "pressure feed" system, filling the vacated parts of the tank with a compressed gas to force the propellants out. This is a very cheap approach, but a pressure-fed rocket engine has a relatively low fuel flow rate, which limits thrust. It also requires that the fuel tanks be built to withstand pressurization, which increases weight. Despite these drawbacks, the simplicity and potential low cost of pressure-fed engines makes them attractive, and there has been work in recent years on pressure-fed engines fueled by LOX and propane. LOX tends to be self-pressurizing as it boils off, and propane becomes self-pressurizing if heated slightly.

Most modern rocket engines use high-speed turbine pumps to drive the propellants into the combustion chamber at a very fast rate. The turbopumps are initially spun up on with a "starter" -- an electrically activated solid-fuel charge. However, even rocket vehicles that use turbopumps still pressurize the propellant tanks, not so much to drive propellant flow rate as to prevent a void from arising as the tank is vacated, which would otherwise stifle thrust or even cause the propellant tanks to collapse.

The fuel line is generally wrapped around the engine nozzle, helping to cool the engine and also "preheat" the fuel, conserving some of the combustion energy that would otherwise be wasted. This is known as "regenerative" cooling. While regenerative cooling is common now, early rockets used water cooling, or various types of "heat sinks" -- sets of fins or other structures that could radiate the heat away.

There is also usually a second set of valves downstream from the turbopumps. These are kept closed until propellant pressures build up enough to ensure a regular fuel flow into the combustion chamber. If the propellants were allowed to trickle unevenly into the combustion chamber, the burn would be irregular and poorly controlled. In other terms, they're just "hold-off" valves, ensuring a crisp start to combustion.

The propellant feed system is also usually designed to make sure that oxidizer flow reaches its proper flow level before fuel does. However, modern liquid fuel rocket engines are often designed to burn "fuel rich", throwing so much fuel into the combustion chamber that all of it can't be burned, meaning limiting the ratio of oxidizer to fuel. That may seem inefficient, but it allows an increase in exhaust mass flow, and so thrust, without raising exhaust temperature. Determining the proper propellant ratio is a fine art.

The propellants are forced into the combustion chamber using an "injector" system that ensures they mix properly, which can be visualized as something like a shower spray head. The engine burn is then initiated with an "igniter" system. Igniter systems can be based on electrical spark plugs; pyrotechnic charges; electroresistive heating elements; and even small igniter rockets using hypergolic storable propellants.

The exhaust then pours out the engine nozzle, producing thrust. The input to the nozzle is a narrow constriction, where the exhaust flows at high pressure and low velocity. As the exhaust flows down the nozzle it expands, losing pressure but gaining velocity, resulting in greater momentum transfer to the rocket vehicle.

There is a trade-off between nozzle size and thrust. If the nozzle is too small, the exhaust will be robbed of thrust. If the nozzle is too big, the exhaust will separate from the nozzle wall, not only making the larger size useless but also counterproductive, since it sets up turbulence that robs the exhaust flow of thrust. The behavior of the exhaust varies with atmospheric pressure, which of course falls off as the rocket climbs, and so optimizing the exhaust design is a troublesome matter. It is also obvious that building a big engine is more difficult and expensive than a small one. As a compromise, rocket engines were developed that featured multiple clustered "thrust chambers" but shared the same systems upstream.

* Usually the interval between spin-up of the starter motor and engine ignition is about a second. Once the propellant is gated into the main combustion chamber, a small amount of it is fed back towards the turbopump drive turbine and burned in a small secondary combustion chamber to take over from the starter, which then burns out. The secondary combustion chamber is primarily intended to drive the turbopump system -- though there are some designs that use preheated fuel, instead of a secondary combustion chamber, for turbopump drive.

The secondary combustion chamber burn must be precisely throttled to ensure proper turbopump RPM. Liquid fuel rocket engines that have variable thrust have adjustable valves on these feed lines, while those that have constant thrust simply use a small orifice to limit fuel rate. Incidentally, the secondary combustion chamber may need an igniter of its own, though in many cases the starter exhaust into the drive turbine will be hot enough to ignite the liquid propellants.

To get really elaborate, exhaust from the secondary combustion chamber has several other uses:

* It is possible to simply let a liquid rocket engine burn until it runs out of fuel, but this causes thrust to "sputter" at the end, and the turbopumps may spin up when they start pushing vapor instead of than liquid, eventually tearing themselves apart. Neither behavior provides a very smooth ride. A better way is to cut off the propellant flow to the secondary combustion chamber driving the turbopumps. Once the turbopumps spin down enough, the valves controlling the fuel flow to the main combustion chamber shut off propellant flow, and the engine burn halts immediately.

Of course, a controlled shutdown is absolutely essential in a restartable engine. As mentioned, ensuring fuel flow in a restartable engine is tricky because the propellants cannot flow down to the engine, since "down" doesn't exist under weightless conditions. There are several approaches:

* Modern LOX-hydrocarbon rocket engines burn LOX and kerosene, or more precisely a highly refined grade of kerosene named "RP-1" or just "RP" for short, which some sources claim stands for "rocket propellant" and others claim stands for "refined petroleum" -- pick one. RP looks and smells like ordinary kerosene, but it is pure and has highly predictable burn and density characteristics. If a large rocket vehicle were fueled with normal kerosene it would burn, but not evenly, and the fuel weight could vary by a matter of tonnes, which would make the performance of the vehicle very unpredictable.

LOX-RP is traditionally one of the most popular liquid-fuel schemes. LOX-RP engines typically have a specific impulse of about 260 seconds. This will be used as a "baseline" value in the rest of this document, with specific impulse values given relative to LOX-RP. This avoids the question of metric or English units, and the less-than-intuitive use of "seconds" as a unit of measure for rocket efficiency.

LOX-LH2 provides higher specific impulse, about 1.5 times that of LOX-RP, but there is a price for it. LH2 is a low-density fuel, meaning it requires a large tank, and so the rocket vehicle has to be big as well. It also has to be kept very cold, substantially colder than liquid oxygen. The LOX and LH2 tanks have to be thermally isolated from each other, or the LOX will tend to freeze and the LH2 to boil.

LOX-kerosene and LOX-LH2 are both "cryogenic" propulsion systems, meaning they require cooled propellants. Both use cooled LOX oxidizer, while LOX-LH2 also requires even colder LH2 fuel. Incidentally, some sources refer only to LOX-LH2 as "cryogenic" propulsion, but this usage seems a bit misleading and is not employed here. A rocket vehicle using cryogenic propellants has to be fueled a relatively short time before launch, or the propellants will gradually vaporize away. Launch systems for such vehicles cycle the cryogenic propellants through an external cooling system to keep losses as low as possible, and also include a bleed valve to keep vaporized propellants from building up in the tanks and possibly rupturing them.

Storable propellants, as the name implies, avoid this problem. They can be loaded into a rocket vehicle and left indefinitely. One of the earliest storable propellant combinations was concentrated hydrogen peroxide (HO), known as "high-test peroxide (HTP)", for oxidizer, and aniline (C6H7N1), a benzene derivative, for fuel.

Since HTP tends to degrade slowly in storage over time, it was generally replaced as an oxidizer by nitric acid (NHO3), nitrogen tetroxide (N2O4), or combinations of the two. Aniline was replaced by hydrazine (NH2NH2); unsymmetrical dimethyl hydrazine (UDMH), with the chemical formula N(CH3)2NH2; or a mix of the two, sometimes called "Aerozine-50". Nitrogen tetroxide and UDMH are a popular combination, and have a specific impulse only a few percent less than that of LOX-RP. Launches of boosters using storable propellants generate a distinctive nasty-looking reddish exhaust cloud.

Storables have a number of drawbacks. Not only are they usually hypergolic, they are also as a rule extremely toxic and corrosive. Storable propellant tanks have to be lined with stainless steel, and workers handling these propellants must wear protective clothing and respirators. There were cases in both the Soviet and American space programs where space capsules returning to Earth fired thrusters fueled by storables too persistently, and the fumes overcame the crews, though nobody was ever done any long-term harm. The US space shuttle's orbital propulsion system used storables, and when the orbiter landed, a crew in protective clothing had to go out and "safe" it before anyone else was allowed to get near. The unpleasant behavior of storables means higher cost, and they are also now generally regarded as environmentally unacceptable and avoided when possible.

Storables were initially used in military missiles that had to be ready to fire on a moment's notice. Their tanks were permanently fueled, capped by seals that were blown at the moment of firing. Solid propellants are now preferred for military missiles, though the Soviets stayed with storables for much longer than the US since they were comfortable with the technology. The Soviets even used them for submarine-launched missiles, despite the threat posed by such toxic and violent chemicals in a closed environment. The successor Russian government has found disposal of large quantities of toxic fuels to be a major environmental headache.

* Storables are still used on spacecraft, such as deep-space probes that will fly through space for years and make occasional engine burns. Obviously there is no practical way to store cryogenic propellants for such a long period of time.

Of course, the thruster systems on such spacecraft are also based on storable fuels. Thruster systems are in general simple, very low thrust, restartable rockets. The most elaborate use storable propellants, typically N2O4-UDMH, but "monopropellant" thrusters are also used. These feature a fuel, usually hydrazine, that burns when passed over a catalyst. Monopropellant thrusters are much simpler and, in principle, more reliable than a bipropellant system, but they are much less efficient, with a monopropellant hydrazine thruster having a specific impulse less than 70% that of LOX-RP. Whether bipropellant or monopropellant, the thrusters are fed using a pressure or electric pump system, as the high propellant flow rates of a turbopump are not required.

Another scheme occasionally used is the "cold gas" thruster, which is nothing more than a compressed-gas jet. Cold gas thrusters are not efficient at all, but they are extremely simple and very safe, since no combustibles or combustion is involved. There are a range of more exotic thruster schemes, discussed later.

* A wide range of different liquid propellant combinations have been used for liquid rocket engines, and the combinations listed above are only those that actually went into widespread use. Unusual fuels include ammonia and ethanol (grain alcohol), and there have been some odd combinations of more conventional propellants, such as LOX-UDMH or HTP-RP.

Liquid diatomic fluorine (LF2) has been considered as an oxidizer in place of LOX, and LF2-LH2 propulsion actually has specific a few percentage points higher than that of LOX-LH2. It has never been used operationally, likely because fluorine is such a nasty substance to handle.

In recent years, tests have been performed of LOX-liquid methane engines. Such a propellant combination is at least competitive with other liquid-fuel combinations, burns cleaner than all but LOX-LH2, and might come in very handy for a Mars expedition, since both LOX and methane could be obtained from the Martian environment -- allowing the crew to "tank up" there before coming back home.

Incidentally, liquid fuels may include small amounts of additives to make them burn more smoothly, make them easier to handle, or keep them from freezing up in space.

BACK_TO_TOP

[1.3] LIQUID PROPELLANT ENGINES: A SURVEY / AEROSPIKE & RBCC ENGINES

* Attempting to provide a comprehensive catalog of rocket engines used on space launch vehicles and spacecraft would be equivalent to cataloging all such vehicles, which would be a fine thing in itself but too painful to deal with here. As far as this document is concerned, a quick sketch of representative or important liquid-fuel engines will be enough.

The first liquid-fuel rocket engine to be produced in quantity was the engine for the German V-2 missile, which burned LOX and ethanol and produced 245 kN (25,000 kgp / 55,000 pounds) of thrust, which was far beyond the capabilities of any other rocket engine available at the time.

V-2 engine

After the war, however, substantially more powerful engines were required, In the early 1950s, the US Air Force (USAF) worked on the development of a supersonic ramjet-powered cruise missile named the "Navaho" that was launched with a liquid-fuel booster. The Navaho program was canceled, but it provided a number of useful technologies for better missiles, particularly a LOX-RP engine that produced about 608 kN (62,000 kgp / 137,000 lbf) thrust, the Rocketdyne "XLR-83-NA-1". This engine was the basis for several of America's early missiles and space launchers.

The Atlas intercontinental ballistic missile (ICBM) and booster was powered by three engines based on the Navaho engine. The half-stage featured two "LR-89" series engines -- which confusingly shared a turbopump, and so could just as well be regarded as a single engine with two thrust chambers -- and an "LR-105" series engine. The entire engine assembly went through a series of designations, from the "MA-1" of flight test prototypes to the "MA-5" that was used in maturity. Liftoff thrust of the full MA-5 engine assembly was about 1,920 kN (195,500 kgp / 431,000 lbf), with the thrust falling to about a third of that when the half-stage was discarded. The Atlas would never be very useful as a weapon, but it would prove a highly successful space launch vehicle.

The "Thor" medium-range missile, which would evolve into the "Delta" series of launch vehicles, was powered by a single "LR-79" or "MB-3" engine, also with 756 kN (77,000 kgp / 170,000 lbf) thrust and based on the Navaho engine.

The initial version of the "Titan" missile was powered by a different type of engine, the Aerojet "LR-87-3", which was a single engine with twin thrust chambers, providing a total of about 1,470 kN (150,000 kgp / 330,000 lbf) thrust, burning LOX and kerosene. The Titan upper stage was powered by an Aerojet "LR-91-3" engine, similar to the LR-87-3 but with a single thrust chamber and providing 356 kN (36,300 kgp / 80,000 lbf) thrust. Later versions of the Titan used modified versions of the LR-87 and LR-91 that burned storable propellants.

Titan RL-87 engines

In the late 1950s, the US wanted to go to still more powerful LOX-RP engines. This led to the development of the Rocketdyne "H-1" engine, with eight used on the first stage of the early "Saturn I" and "Saturn IB" boosters, providing 837 kN (85,300 kgp / 90,000 lbf) thrust each. An improved version of the H-1 designated the "RS-27" was used for later variants of the Delta launch vehicle. The H-1 also was the basis for the scaled-up Rocketdyne "F-1" engine, with 6,671 kN (680,000 kgp / 1.5 million lbf) thrust. Five such engines powered the first stage of the "Saturn V" booster that sent Americans to the Moon.

In the meantime, the US had been pioneering LOX-LH2 propulsion, first developing the Pratt & Whitney "RL10" engines with 66.7 kN (6,800 kgp / 15,000 lbf) thrust each, with two powering the "Centaur" upper stage of the "Atlas Centaur" booster. The Centaur upper stage still retains this engine in the 21st century, in the form of the "RL10A-4" variant, and also has been used on the upper stages of the Boeing "Delta III" and "Delta IV" boosters, in the form of the more advanced "RL10B-2" engine. A throttleable version, the "RL10A-5", was used on the McDonnell Douglas DC-X/A demonstrator spacecraft in the 1990s.

The next step from the RL10 was the much more powerful "J-2" LOX-LH2 engine, providing 105.95 kN (108,000 kgp / 238,000 lbf) thrust. NASA is now planning a simplified and uprated version of the J-2, the "J-2S", with 1,179 kN (120,200 kgp / 265,000 lbf) thrust, for the new crewed Moon effort.

Rocketdyne J-2 engine

The J-2 led to the Rocketdyne "Space Shuttle Main Engine (SSME)", which as its name states was the primary powerplant of the NASA space shuttle. The shuttle used three SSMEs, with 2,090 kN (213,000 kgp / 470,000 lbf) thrust each. The SSME's development was notoriously troublesome, since it was pushing the state of the art and the program was underfunded. It did become a reliable and effective engine, and a simplified non-reusable derivative may be used on a future US heavy-lift booster.

The Rocketdyne "RS-68" engine, used on the Delta 4 series of boosters, is currently the world's biggest LOX-LH2 engine, with a height of 5.18 meters (17 feet), a nozzle diameter of 2.44 meters (8 feet), and a launch thrust of 2,918 kN (297,500 kgp / 656,000 lbf). The RS-68 is an entirely new design, but incorporates lessons learned with the SSME. The RS-68 took only half as long to develop, has only 20% of the number of parts of the SSME, and is much less labor-intensive to build.

The most recent big US rocket motor is the "Merlin 1", developed by the SpaceX commercial spaceflight firm for that company's Falcon boosters. The Merlins are turbopump-driven, burning LOX-kerosene; they are innovative in that they use a carbon-composite nozzle. The initial "Merlin 1A" variant provided 340 kN (34,920 kgp / 77,000 lbf) thrust at sea level; it evolved to the "Merlin 1D" variant, with 650 kN (66,650 kgp / 147,000 lbf) sea-level thrust.

* The primary Soviet rocket engines in the early days of the Space Race were the "RD-107" and "RD-108". These were basically "sister" designs, each with four thrust chambers and a single turbopump, differing in that the RD-107 had two vernier engines, while the RD-108 had four. They both produced 907 kN (92,500 kgp / 204,000 lbf) of thrust, and were used on the SL-1 launch vehicle that put Sputnik 1, the first Earth satellite, into orbit, which ultimately evolved into the modern "Soyuz" booster. The SL-1 featured a "core" first stage with an RD-108 engine, surrounded by four liquid-fuel boosters, each with an RD-107 engine.

RD-107 engine

The Soviets followed the SL-1 series of boosters with the much more powerful "Proton" booster, which put the "Salyut" and "Mir" space stations into orbit. The first stage of the Proton is powered by six "RD-253" engines using N2O4 / UDMH storable propellants and providing 3,090 kN (315,000 kgp / 695,000 lbf) thrust each.

Later, the Soviets developed a much more powerful LOX-RP four-chamber engine, the "RD-170", with 8,182 kN (834,000 kgp / 1.84 million lbf) thrust. The first stage of the "Zenit" launch vehicle is powered by a single RD-170 engine. The Soviets also developed a variant of the RD-170 designated the "RD-180" with two thrust chambers instead of four and, unsurprisingly, half the thrust.

In an interesting irony, the RD-180 was adopted for the modern variants of the Atlas booster, the "Atlas III" and the "Atlas V". Both these launch vehicles abandoned the old half-stage scheme and used a single RD-180 engine. The RD-180 engines were provided for these boosters by a collaboration of NPO Energomash of Russia and Pratt & Whitney in the US. Given poor relations between the US and Russia as of late, there's been a push to replace the RD-180 with a US design.

Of course, liquid-fuel rocket motors were developed by other nations besides the US and USSR, but it should be no slight to say that, having covered American and Soviet engines, all the rest don't bring much more to the party, and further discussion of engine types would become tedious.

* As far as storable propellant engines used on spacecraft themselves go, one of the classic examples was the "Service Propulsion System (SPS)" for the Apollo Command & Service Module (CSM). The SPS generated 91.2 kN (9,300 kg / 20,500 lb) of thrust. It was a fully restartable pressure-fed engine, with no turbopumps, featuring redundant subsystems where possible; its design philosophy was to make it as simple and reliable as possible. Of course, the Apollo Lunar Module (LM) also used storable propellant engines, with the descent stage engine providing 44.5 kN (4,535 kg / 10,000 lb) thrust and the ascent stage engine providing 15.6 kN (1,590 kg / 3,500 lb) thrust.

As mentioned, the space shuttle orbiter had a secondary set of storable propellant engines, known as the "Orbital Maneuvering System (OMS)", for maneuvers after arrival in space. Each of the two OMS engines on the orbiter provided 27 kN (2,750 kg / 6,070 lb) of thrust.

* While conventional chemical rocket engines continue to be refined, work has also been performed on new configurations, such as the "linear aerospike" and "rocket-based combined cycle" engines.

NASA was working with private industry on an aerospike engine for the canceled X-33 experimental reusable launch vehicle. An aerospike engine is very different in appearance from a conventional rocket engine. Instead of a bell-shaped nozzle, the aerospike looks like a wedge rammed into the rear of a vehicle, with nozzle ports, or "thrust cell chambers", along each side of the base of the wedge. Thrust cell chamber exhaust is confined on one side by the wedge, with air pressure providing confinement on the other side. Propellant pumps and other hardware are contained inside the wedge.

Other rocket companies have continued to experiment with the aerospike concept, launching small rockets that feature a spike surrounded by exhaust holes. As with the canceled X-33 engine, confinement is provided by the spike on the inside and air pressure on the outside. The advantage of the aerospike nozzle is that it automatically adjusts for air pressure, while a conventional bell nozzle is designed basically around one value of air pressure. Aerospike advocates believe they can achieve efficiencies a quarter to a third better than those of a conventional rocket nozzle.

The RBCC engine, as its name more or less hints, is a combination of jet engine and rocket engine. An RBCC engine consists of a "supersonic combustion ramjet (scramjet)" -- basically just a "stovepipe" with an air intake in front, fuel injectors and igniters in the middle, and an exhaust at the back -- but with non-airbreathing rocket nozzles placed inside, within the flow path. Early RBCC engines will use LOX-hydrocarbon propulsion, but LOX-LH2 propulsion is expected over the long run.

An RBCC-powered spacecraft would take off using rocket propulsion, with the airflow through the ramjet duct helping boost thrust through simple momentum transfer. At about Mach 2.5, ramjet propulsion would take over, moving to scramjet mode at about Mach 6. At Mach 8 to 12, the spacecraft would be above the atmosphere, and the RBCC engine would then return to rocket propulsion to lift the craft into space.

BACK_TO_TOP

[1.4] SOLID PROPELLANT ENGINES: FUNDAMENTALS

* Solid-fuel rockets go back for centuries, and their fundamental form has remained much the same. They contain propellant formed in a solid cylindrical block, or "grain", normally with a hole up the center to ensure uniform burning of the propellant. The hole can have a variety of configurations that give different "thrust profiles": for example, a cylindrical hole provides thrust that starts out low and increases as the hole widens, giving more area into the grain. Holes generally have configurations in the shape of various kinds of stars, which give the maximum area and thrust for initial burn. The hole is created by packing the fuel around a central "mandrel" piece that is removed after the packing is complete. A nozzle is generally attached to the base of the grain to provide better control of the exhaust.

The traditional solid propellant, invented by the Chinese in the Middle Ages, is "black powder", a mixture of a fuel composed of charcoal and sulfur, and an oxidizer consisting of "saltpeter" (potassium nitrate / KNO3). Black powder fireworks rockets are still in widespread use, since black powder does that job fine and is cheap. Conceptually similar military rocket projectiles were developed in World War II that used "double base" explosives, a combination of nitroglycerine and nitrocellulose (guncotton), as a propellant instead of black powder.

It is difficult, effectively impossible, to build a large rocket vehicle that uses black powder or double base explosives as a propellant. Such materials are both troublesome and dangerous to handle in large quantities. Another problem is that they burn very rapidly, providing high thrust for a short period of time. This is fine for simple, short range rocket projectiles, but not for putting payloads into space.

The first attempts to develop modern solid propellants were performed during World War II. American researchers developed small rockets that used asphalt as a fuel, mixed with potassium perchlorate (KClO4) as an oxidizer. This propellant provided lower thrust over a longer period of time. Asphalt was a poor binder, however, tending to crack at cold ambient temperatures, with the cracks interfering with the burning process; and to flow at high ambient temperatures, requiring the rockets to be stored nose-down. It also produced a great deal of black smoke, which caused particularly difficulties for use as "rocket assisted take-off (RATO)" boosters for aircraft. If one aircraft took off with RATO, the next following it would have to take off through a black haze that blocked the pilot's field of view.

After the war the Americans moved on to more sophisticated solid propellants, using synthetic polymers, particularly synthetic rubbers such as butadiene tire rubbers, mixed with ammonium perchlorate (NH4ClO4) oxidizer (which provides higher performance than potassium perchlorate and burns cleaner) and large proportions of powdered aluminum. The powdered aluminum burned at a high temperature, helping improve thrust. Use of high proportions of aluminum had been held up for some time because the conventional wisdom said that proportions greater than 5% wouldn't burn, but that turned out to be superstition when researchers ignorant of this "rule" tried higher proportions and found they got unprecedented levels of thrust. Later on, small proportions of iron oxide were included to provide a high-temperature "thermite reaction" with the aluminum powder. Modern solid fuels provide a specific impulse only a few percentage points below that of LOX-RP.

The fact that the propellant mix was based on rubberlike polymers allowed the "grain", or solid-rocket fuel element excluding the nozzle), to be cast in large blocks that resisted shrinkage or cracking, which would have affected the continuity of their burn at the very least and caused catastrophic failure at worst.

solid rocket grain

Electrical resistance heating wires could be inserted in the bottom of the grain as igniters. Since a safe solid fuel had a high ignition temperature, the wires sometimes ignited a primer element that had a lower ignition temperature, with this primer then igniting the grain. The primer might even have two "stages", with one element with a low ignition temperature igniting one with an intermediate ignition temperature, with the second element then igniting the grain.

The central mandrel that defined the hole configuration for the desired flight thrust profile was coated with Teflon polymer to allow it to be removed from the cured grain. In the early days of solid fuel rockets, grease was used instead of Teflon, but the grease contaminated the solid fuel and that approach was abandoned.

Modern solid-fuel rocket engines are ideal for military applications. They can be stored almost indefinitely and aren't overly fussy about how they are handled, and they can be used and launched at any time with little preparation. Solid fuels are also denser than liquid fuels, allowing missiles to be more compact, if not that much lighter. This compactness was a particular plus for the development of solid-fuel long-range strategic nuclear missiles, since it allowed them to be stored in a smaller and cheaper missile silo, or be carried on a submarine.

Considerable effort was invested in the late 1950s to develop processes to manufacture the large grains for the "Minuteman" ICBM and the "Polaris" submarine-launched ballistic missile (SLBM). These processes were not trivial: the ammonium perchlorate oxidizer had to be very finely and uniformly milled; all the materials in the solid fuel recipe had to very uniformly mixed; and the grains had to be poured and then cured for several days in a vacuum environment to prevent bubbles from forming.

Work was also done to synthesize high-performance solid fuels, which included proportions of nitroglycerine and a particulate form of nitrocellulose, and in some cases a high explosive known as HMX as well. These high-performance mixes were unsurprisingly less stable and more troublesome in every respect than conventional mixes, and so they were only used for small final stages that were more easily handled.

The work on developing long-range solid-fuel missiles also led to technology for solid-rocket flight control, and for "thrust termination", or shutting down the engine of the upper stage so the warhead could separate and proceed on its proper trajectory to its target.

Small solid-fuel missiles can use fins for flight control, but that isn't practical for long-range missiles since they boost out of the atmosphere, making fins so much dead weight. A steerable nozzle can be used as with a liquid fuel engine, but a simpler scheme was adopted for the Minuteman and Polaris, using a pivoting ring around the lip of a fixed nozzle to redirect the thrust. The ring was known as a "jetavator". Later, inert freon gas was selectively injected into the throat of a fixed nozzle to deflect thrust. As far as thrust termination went, the problem was that once a solid rocket motor is lit, it burns to termination, and there's no way to shut it off. There is a way to cheat, however, by venting the thrust from the stage to the sides or forward so the warhead could continue on its way by itself.

These missile projects also led to the use of fiberglass for motor casings, reducing the weight of the rocket motor and so increasing payload. Building large "filament wound" casings was troublesome, however, and so most casings were made of metal, usually steel, though lightweight titanium could be used when weight was an issue and cost not too painful an issue.

The size of solid-fuel grains grew by bounds through the 1950s, and by the early 1960s they were so big that they were becoming too bulky and heavy to handle and transport in any sensible way. To get around this problem of scale, a new technology, "segmentation", was developed in which the grains were fabricated as cylindrical segments, with the segments locked together in a single solid rocket motor using "lock rings". By the mid-1960s, segmented solid-rocket boosters (SRBs) were being manufactured that could provide heavy thrust to help put increasingly large payloads into orbit.

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[1.5] SOLID PROPELLANT ENGINES: A SURVEY / NEXT-GENERATION SOLIDS

* As with liquid-fuel rocket engines, trying to write a detailed history of solid-fuel rocket engines here would be impractical, and a short survey will have to do. The first modern solid-fuel rockets were developed in the late 1940s, for use as RATO boosters and to power relatively small missiles, such as air-to-air missiles (AAMs). One particularly important set of examples of this period were the "T-41" and "T-42" solid rocket motors for the Hughes Falcon AAM, which would lead to the "Star" motor, used on the Thor Able-Star booster of the late 1950s as a final "kick" stage. A wide range of solid-fuel kick stages would be developed over the following decades.

small solid rocket motors

The first large solid-fuel grain to be developed was the Thiokol "RV-A-10", with a diameter of 79 centimeters (31 inches) and a length of 4.37 meters (14 feet 4 inches). A refined variant was used for the Sergeant battlefield missile. There had been doubts up to that time that it was practical to built large solid-fuel rocket grains, but the RV-A-10 proved beyond doubt that it was possible.

The RV-A-10 led the way to the much larger grains for the three-stage Minuteman ICBM and the Polaris SLBM. The solid-rocket motor for the first stage of the Minuteman was an order of magnitude bigger than the RV-A-10, being 1.67 meters (5 feet 5.7 inches) in diameter and 7.42 meters (24 feet 4 inches) long. Both the Minuteman and Polaris had filament-wound final stages, though the first large-scale filament-wound stage had been flown on a Vanguard satellite launch vehicle in 1959. High-performance solid fuels were also introduced for the Minuteman and Polaris in later versions, and a later version of the Minuteman would also have a second stage with a titanium casing. Ironically, the US imported most of its titanium from the Soviet Union. Since there were very few civilian uses for titanium at the time, the Soviets had to be aware that it was mostly being used in American weapon systems that threatened the USSR.

Work on large solid-fuel grains for the Minuteman and Polaris programs also led to the development of the first all-solid-fuel space launch vehicle, the LTV "Scout", which would have a long career putting small payloads into space. In addition, the Minuteman development effort had a direct connection to the development of segmented solid rocket motors, with Aerojet testing the concept in 1961 by the simple measures of cutting a Minuteman first stage in half and then splicing it back together with a lock ring joint.

Both Aerojet and United Research, which would later become a component of the later United Technology Center (UTC), performed further work and static test firings of segmented rocket boosters. In 1962, UTC won a contract from the USAF to build a five-segment SRB for the Titan III space launch vehicle. The resulting SRBs went into service in 1965, with two SRBs straddling the liquid-fuel Titan core. The initial Titan III SRB motor was 3.05 meters (10 feet) in diameter and 25.8 meters (84 feet 8 inches) long. It led in the 1980s to the 5.5 segment motor for the Titan 34D and the subsequent 7 segment motor for the Titan IV, which produces about 7,565 kN (771,000 kgp / 1.7 million lbf) thrust per SRB.

The biggest solid-rocket motor ever to be put into operation was the SRB for the US space shuttle. Each SRB was 45.5 meters (149 feet 2 inches) long. The precise composition of the shuttle SRB grain by weight was:

The shuttle SRBs were made up of four segments stacked on top of each other. The hole up the center of the SRBs is cone-shaped at the bottom, leading to an 11-point star that runs to the top. This scheme gave maximum thrust of 11,770 kN (1.2 million kgp / 2.65 million lbf) at liftoff, falling off to a sustained level of thrust after that. The nozzle was steerable.

shuttle solid rocket booster

In the US, solid rocket booster technology for spaceflight is now effectively monopolized by Orbital-ATK, the consolidated successor to earlier US solid-rocket companies. It continues to sell a range of different solid-rocket motors:

* As mentioned, one of the major weaknesses of the solid-fuel rocket is the fact that, once lit, it burns to completion, and the only thing that can be done is to divert the thrust when it is no longer needed. The lack of burn control for solid-fuel rockets has led to the development of "hybrid" rockets that use a solid-fuel core along with a liquid oxidizer. The solid fuel component in a hybrid rocket is not impregnated with large quantities of an oxidizer material, which makes the rocket much safer to handle and store since it cannot burn efficiently on its own.

Lockheed Martin has static-tested a hybrid motor with a butadiene-type solid fuel and liquid oxygen oxidizer. Lockheed Martin has also investigated the use of paraffins as propellants; "paraffins" in this case of course refers to the American usage of the term, meaning candle waxes and related solid hydrocarbons, and not the British usage of the term, which is what Americans call kerosene. Purdue University has performed small-scale experiments with hybrid rockets using a butadiene-based solid with storable hydrogen peroxide oxidizer.

Burt Rutan's commercial suborbital crewed spacecraft, "Spaceship One", used a hybrid propulsion system, with a butadiene-type solid fuel and nitrous oxide oxidizer. In this case, the propulsion system is designed for low cost and ease of handling instead of optimal thrust levels. Spaceship One is probably the first thing resembling an operational space vehicle to use hybrid propulsion, and after many years of tinkering the technology seems to be coming of age.

* Experiments have also been performed on another approach to the same problem, in the form of "propellant gels". The idea is to take storable propellants and turn them into gels: hydrazine can be gelled by adding cellulose, and nitric acid can be gelled by adding silicon dioxide (sand, more or less). The results have the consistency of toothpaste. Aluminum powder can be added to provide more "kick".

Since hydrazine and nitric acid are hypergolic, if the two gels come in contact with each other they burn spontaneously -- but not for long, since a crust builds up between them that inhibits further combustion. This makes gels much safer to handle than their liquid forms. To get them to burn in a combustion chamber, they are fed under pressure through an orifice that turns them into an aerosol, allowing them to mix properly. The potential advantages of this approach are high energy density, throttleable operation, and relative safety in handling. Experiments have been performed in determining the suitability of gelled propellants for military missiles. The primary difficulty is that gelled propellants are more expensive than solid fuels.

* One of the more surprising schemes for solid fuel propulsion involved mixing nanoscale aluminum powder with water, and then freezing the mix into a slush. Surprisingly, the aluminum-ice ("ALICE") propellant burns and burns well, with the water combining with the aluminum nanopowder to form aluminum oxides and diatomic oxygen. The scheme was initially test-flown in 2009; work is being done towards additives that would make the fuel more efficient. The ALICE propellant is seen as interesting for "in-situ resource utilization", with a probe taking the aluminum powder to, say, Mars, and then mixing it with local water to boost a sample back to Earth.

Another interesting concept in solid fuel propulsion, devised by a British-Ukrainian team, is the "autophage" rocket -- which, instead of staging, simply consumes itself as it ascends, becoming shorter and shorter. The grain would consist of a propylene cylinder, presumably mixed with powdered aluminum, that would provide fuel and act as an airframe, with powdered oxidizer in the center, made of ammonium perchlorate and ammonium nitrate.

The grain itself would not burn. It would sit on top of a "vaporization surface", which is electrically heated to vaporize the propellants before they enter the combustion chamber of a rocket engine. As the rocket ascends, inertia forces the grain against the combustion surface, slowly consuming the grain. The rate of feed of the grain can be controlled, allowing the rocket to be throttled. The autophage scheme is not seen as very useful for large rockets, but it could mean cheap small ones. Nobody's flown an operational autophage rocket yet.

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