Since the beginning of the space age, liquid-fueled rockets have provided the basis of everything we have accomplished in space. Whether it has been the mighty F-1, which powered Apollo to the Moon, or the workhorse RL-10, or the high performance SSME, all of the major missions depend on liquid-fueled propulsion. And as we move forward to the next generation systems, the success or failure of these systems will be intimately tied to the success of its liquid-fuel propulsion system.
The following discussion focuses on the design of liquid-fueled rocket systems and examines the design choices that must be considered in integrating the propulsion system for a selected application. It is to be understood that liquid-fueled systems refer to the oxidizer as well as the fuel and that both of these are carried aboard the vehicle. Additional detailed information on the design of liquid propellant rocket engines can be found in References 1-3.
Today’s liquid-fueled rocket propulsion designs originate from the theories of a Russian, Konstain E. Tsiolkovsky, and from experiments performed by an American, Robert H. Goddard and a German, Herman Oberth. Goddard and Oberth made some of the first working liquid propellant rocket engines. Goddard was first to successfully test a liquid-fueled engine, which he did in 1926. Because of lack of interest by the United States government, Goddard worked primarily as a loner with a small cadre of technicians to support his work in the White Sands, New Mexico desert. He contributed to U.S. rocket development during World War II, but the importance of his contribution to rocket development was not recognized until the beginning of the space age. Herman Oberth, working with young Werner von Braun, developed liquid propellant rockets that got the attention of the German military. Their efforts culminated in the development of the military A-4 (or V-2) rocket. After World War II, Werner von Braun and many expatriated German scientists came to the United States. Their expertise and technical capabilities led to the development of the Jupiter and
Redstone missiles and the Saturn series of manned launch vehicles for the Apollo Program.
Rocket Propulsion Systems
A liquid-fueled rocket propulsion system consists of a number of carefully integrated components that must be carefully chosen and designed to function as part of the whole. The operation of an integrated system can best be viewed by identifying the required functions that must be satisfied and then considering the components that are chosen to fulfill these functions. The primary function of a rocket propulsion system is to produce thrust that accelerates the spacecraft to the final velocity required to achieve its goal. During the flight, the thrust must be controlled in magnitude and direction, allowing the space vehicle to be accelerated and guided along a desired flight trajectory. To produce thrust, the propellants are burned in a combustion chamber at high pressure, yielding high temperature gases that are subsequently expanded and accelerated through a convergent-divergent nozzle and ejected at high supersonic velocity. Thrust is produced in the direction opposite to the mean flow of the ejected hot gas at a magnitude equal to the difference in momentum between the incoming liquid propellants and the ejected hot gas plus the difference between the product of the pressure and the area at the nozzle exit and chamber.
To produce a practical propulsion system, the integrated system must control the flow of propellants from the storage tanks, deliver them at the required pressures, and inject them into the combustor. This process can be achieved either by pumping the propellant to the required pressures (pump-fed system) or by pressurizing them with high-pressure gas in the storage tanks (pressurized system). The integrated pumped system must provide the power required to drive the pumps and provide cooling for the thrust chamber (i.e., combustor chamber and nozzle) surfaces exposed to the hot combustion gases. For pressure-fed systems, a gas pressurization system must be provided. The nature of the component requirements and the relevant design options depend on the application for which the system is being designed. The design process for the propulsion system must consider the following questions:
• What is the spacecraft application?
• What are the vehicle’s operating requirements?
• Which fuel and oxidizer should be used?
• Which type of propellant supply system?
• What type of propellant pressurization?
• What provides the engine pumping power?
• Which type of thrust chamber design?
Part of this article is structured around the subsystems of an engine. But it is important to keep in mind that these engines are highly coupled systems, from both the viewpoint of mechanical systems and fluid flow. For instance, in a pre-burner cycle, the size of the pumps defines the fraction of the propellants used to drive the turbopumps, and this combined with the selected preburner propellant mixture ratio defines the maximum combustion chamber pressure. The preburner mixture ratio defines the turbine operating temperature and the potential need for high temperature materials, or in maximum performance designs, the need for cooled turbine components. Thus, no design of a subsystem is really complete until its interactions with the rest of the system are understood.
An overview of some of the design issues would be helpful to put things in perspective. Why consider the spacecraft application during engine design? Because requirements arising from the application become a strong driver for defining the selection of propellants and the design features of the integrated propulsion system. Should the system be designed to perform best in the atmosphere or in vacuum conditions? Are refrigerated cryogenic propellants acceptable, or is long-term storage important in the application? What level of engine combustion pressure should be considered, and what range of throttle control is important? Power to drive the pumps must be generated by the engine cycle. Which methods are available, and what is the best choice for a selected application? What issues arise from the fuel tanks and the associated fuel delivery system? Does the engine run only at launch time, or will it be used again after coasting in space, and how does this issue affect the design choices? What about the design of the combustion chamber and thrust nozzle (thrust chamber)? Which design features should be considered to yield the required engine performance and to provide adequate cooling of the structure? What considerations should be made with regard to manufacturing capabilities during the design of the engine? These issues are addressed in the sections that follow.
Definition of Designs. The following design parameters have determined the selection of specific liquid rocket propulsion on past and present launch vehicles and spacecraft:
1. Propellant chemistry. Propellant chemistry is selected to provide the highest performance for a feasible mechanical design that is safe to operate and is consistent with storage requirements.
2. Vehicle performance. Vehicle performance is maximized by optimizing the propulsion design to yield the maximum performance at the lowest weight and with the smallest geometric package to interface with the vehicle. When maximizing vehicle performance, optimization of the propellant combination affects the vehicle in two ways: (1) the propellant bulk density that affects vehicle volume and (2) the specific impulse (Isp) that affects vehicle gross weight. Bulk density is the ratio of the total mass of propel-lants burned to their total stored volume. Isp is the thrust produced per unit mass of propellants burned. To produce a specific amount of vehicle acceleration, higher Isp reduces the mass of propellants required, and higher bulk density reduces the size of the propellant storage tanks.
3. Operating environment. Three major application categories exist in designing liquid rockets: boost from sea level, boost in near-vacuum or vacuum conditions, and in-orbit operations. The last category may be further divided into two subcategories, orbital maneuvering and station keeping.
4. Cost versus performance. Each engine design represents a compromise between high performance and affordability. The mission requirements and the financial realities for the planned vehicle must be considered when selecting a system design.
5. Payload. The question of essential importance here is whether or not the vehicle will be carrying passengers. It will be shown that the requirements of human-rating flight hardware have driven engine choices and affected performance, cost, and reliability.
6. Reusability of the rocket stages. Whether the hardware is to be used again or is part of an expendable system affects the choice of engine type and also the complexity of the engines.
Items 3 through 6 will be discussed immediately followed by a more in-depth discussion of 1 and 2.
Operating Environment. The requirements for launch-vehicle boost from sea level have traditionally driven first-stage systems toward high-thrust, moderate-performance engines. Kerosene has traditionally been the most common fuel for first-stage propulsion systems, and liquid oxygen is the oxidizer. The primary advantages of this choice are ease of propellant handling because kerosene is storable at ambient temperatures, and high propellant bulk density. Sea level specific impulse Isp (see Eq. 5) from 250-300 seconds is typical of such engines. This is not as high as it would be with an oxygen-hydrogen system, but performance is not as crucial for first stages as it is for upper stages. Weight increase in the first stage does not necessarily force resizing the other stages. In addition, the higher propellant bulk density allows the construction of smaller fuel tanks. Because the first stage is typically the largest for any launch vehicle, decreasing its size reduces the cross-sectional area of the whole vehicle and reduces drag losses. An example of a kerosene-burning first stage is the SI-C on the Apollo Saturn V vehicle. Five Rocketdyne F-1 engines operating on a gas generator cycle comprised the propulsion for this stage. Nearly every launcher in the U.S. fleet has used Earth-storable fuel in its first stage. The first stage of the Atlas and Delta rockets are kerosene-powered; the Titan rockets use Aerozine-50 (50-50 hydrazine/UDMH) and nitrogen tetroxide as the oxidizer. The Space Shuttle Main Engine (SSME) is the major exception to this. Liquid hydrogen and liquid oxygen (LOX) were chosen as the propellant combination because of the desire to maximize performance to achieve orbit. Note that all four of the rocket families mentioned that have first stage hydrocarbon engines used cryogenic-fueled upper stages.
A large number of upper stages used liquid hydrogen (LH2) and liquid oxygen as propellants. Cryogenic fuels are used in upper stages primarily because of the high specific impulse, which minimizes the overall stage weight needed to create the required change in velocity. Reducing the weight for a vehicle’s second or third stages is highly desirable because the size of the stages below may then be decreased as well. Pioneers such as Tsiolkovsky recognized the utility of liquid hydrogen as a rocket fuel. The development of cryogenic technology accelerated after World War II, pushed by the desire to take advantage of hydrogen’s specific impulse potential. The first liquid-hydrogen-fueled rocket engine to be used on a launch vehicle was the Pratt & Whitney RL10, which debuted on the Atlas-Centaur stage in 1963 (4). The Centaur was also used on the Titan booster, and a new version of the RL10 powers the upper stage of the new Delta III. The Saturn V used Rocketdyne’s hydrogen-fueled J-2 engines for its second and third stages and demonstrated the benefit of cryogenic upper stages on a vehicle that has a storable first stage.
As noted before, the third category of operating conditions is that for in-orbit operations. The need for high reliability and restart capability in propulsion dictates that the engine be as simple as possible. Typically, this requirement has been met with hypergolic propellants. These are fuels that ignite on contact and therefore do not need a separate ignition system. The specific impulse of hyper-golic propellants is not as high as that afforded by hydrogen, but high performance is not crucial for in-orbit propulsion because the change in velocity is usually small. Most propellants used for in-orbit propulsion systems are also storable, a quality made necessary because they may remain in the vehicle’s tanks for an extended period of time before use. A prime example of the use of storable hypergolic propellants for orbital maneuvering is the Space Shuttle’s Orbital Maneuvering System (OMS). The OMS is used to provide the final velocity change necessary to inject the shuttle into its orbit, perform orbital changes during the mission, and provide the burn to reduce orbital velocity for deorbiting. These two engines use monomethylhydrazine and nitrogen tetroxide and develop 6000 pounds of thrust each. In addition to the OMS engines, the Shuttle has a reaction control system that is used to orient the vehicle in space, provide some deorbit velocity change, and separate the Orbiter from the external propellant tank. These two motors are also powered by monomethylhydrazine and nitrogen tetroxide and are pressure-fed, rather than pump-fed, which greatly contributes to their reliability and simplicity (5). Pressure-fed engines were common in boosters in early rockets flown by the Germans and by pioneers such as Robert Goddard, when launchers had only a fraction of the velocity required to achieve orbit. Pressure-fed engines are used for nearly every in-orbit application today because they are optimized for velocity requirements that are not as great as booster stages.
Cost versus Performance. The goal of any rocket development program is to achieve the highest possible performance and reliability within the financial constraints of the program. At no time have the imperatives of cost containment been so important as they are today, a fact reflected in most new engine designs. Total ”cost of ownership” can be defined as the sum of the development cost, cost of procurement of production units, operational cost, and the maintenance cost. The importance of each of these various elements depends on the application requirements and on the operational goals. The cost of development is directly related to the complexity of the system. It can be reduced if systems can be developed as individual components and the developed components are subsequently integrated. In general, engines that have simple power cycles and lower chamber pressure cost less than high-pressure, high-performance systems. Interpreting this statement with regard to the type of engine cycle, gas generator cycles have a lower development cost than preburner cycles because operating pressures can be lower and power cycles can be developed independently of the thrust chamber operation.
The Rocketdyne RS-68, a liquid hydrogen/liquid oxygen engine to be used on the first stage of the Delta IV launch vehicle, is a gas generator cycle engine designed for simplicity, low cost, and moderate performance. The selection of the gas generator cycle is an important part of this cost reduction because the increase in engine simplicity and decrease in chamber pressure typically associated with gas generators reduce the cost of the system.
Alternatively, the expander cycle, used currently on the RL10 engine with liquid oxygen and liquid hydrogen, provides a means to avoid the performance losses inherent in gas generator engines, while providing potential for reduced cost. This cycle has the attributes for very low cost of manufacture and still delivers greater performance than the gas generator cycle.
On the other hand, the Space Shuttle Main Engine was designed to meet much more stringent performance requirements, without being as limited by financial restrictions as modern expendable launch vehicles. The performance requirements dictated the selection of liquid hydrogen as the fuel and staged combustion as the engine cycle. Both choices resulted in greater complexity and cost for the engine but resulted in a vacuum specific impulse in excess of 450 seconds.
It is recognized more and more that the most significant factor in cost is the design process itself. As a result, many new techniques are being used and are achieving great success. The most successful is the use of integrated product teams to execute the design and development process. These are colocated mul-tidisciplinary teams that include every function from design engineers to suppliers. These teams ensure that producibility and inspectability are built into the product when it is being designed for performance, rather than being added later, as is much more costly and was typical of the past. Another example is using advanced computer-aided design models that allow instantaneous sharing of information by all members of the design team.
Finally, when the design team focuses on cost, it leads to decisions that may not be made otherwise. Choices that lead to a more moderate environment result in lower pressures, temperatures, pump speeds, vibrations, and heat loads. All of these have a significant impact in reducing cost.
Payload. Most rocket engines are designed to carry unmanned payloads into space. Although reliability is important in these power plants, it becomes much more crucial for human spaceflight. When lives depend on the hardware, choices are made to maximize the probability that the engine will work correctly. Reliability was the foremost concern of designers of the Apollo propulsion systems. Examples are seen in the Apollo Service Propulsion System and the descent and ascent engines for the Lunar Module. If one of the elements had failed to work, astronauts might have been stranded in orbit around the Moon or on its surface. The first choice made to increase the dependability of the Apollo systems was to make all engines pressure-fed. The absence of pumps eliminated a source of potential failures. The choice of propellants also reflected the desire for reliability. The Service Propulsion System, the engine providing the velocity change necessary to enter and leave orbit around the Moon, used nitrogen tetroxide and a 50/50 mix of hydrazine and unsymmetrical dimethylhydrazine as fuels. This combination is hypergolic, and both propellants are storable. Both engines for the lunar module also used storable, hypergolic propellants (6). Performance requirements for the SSME dictated the propellant choice and cycle choice. Reliability was achieved through rigorous engineering development.
The choices of cycles and propellants are not the only criteria that affect performance, cost, and reliability. The selection of component materials can affect the overall engine design characteristics. One example is the nozzle on the lunar module ascent engine. The choice for this nozzle was to use ablative material that would be thermally eroded as the engine operated. This simple nozzle design that eliminated the need for cooling passages and complicated plumbing was appropriate for an engine which would have to start only once, and yet would have to operate as designed since the first time it was used a long way from Earth. Reusability. It is widely held that the only real means to achieving the low-cost access to space that the country needs is through extensive use of reusable systems. No matter how low the cost of vehicles and engines is driven, these are still highly complex systems and will never be cheap enough to throw away after each flight.
The only current reusable system is the Space Shuttle, although this has not come close to achieving its design goal for reusability. The design specification during the original design studies for the SSME was 100 flights. Important advances in technology have occurred since the SSME was designed. There are vastly improved tools for understanding internal hot gas flow. These are needed to determine the thermal and mechanical loads (both steady and unsteady) and to predict when parts would ultimately fail. There is improved understanding of both low-cycle and high-cycle fatigue. In addition, there are advanced materials that resist these hostile environments better and thus aid in extending service life. Finally, there are new component designs, such as hydrostatic bearings. Future systems will employ many of these technologies whatever the design requirement, but reusable systems will benefit most.
As the increased emphasis on cost drives changes in design philosophy, arguably the greatest benefit to reusability will come from design decisions. Decisions that focus on life-cycle cost rather than only on development cost will lead to different systems. Similar trade-offs must be made with respect to performance. It is extremely unlikely that a highly reusable system will also provide the ultimate in performance. The challenge for today’s propulsion system designer is to provide sufficient performance to achieve the mission while ensuring enough margin so that the system is truly reusable.
Propellants are typically chosen on the basis of several different parameters. In general, these are the density and the molecular weight, the storability (Earth or space ambient temperatures considered), and the ratio of oxidizer to fuel that optimizes the given combination for maximum performance and bulk density. The performance for given propellant combinations is a function of the oxidizer-to-fuel mixture ratio that maximizes the temperature of combustion, which generally maximizes the performance (specific impulse). The bulk density is the relative density of the mixture when considered as a total mass and total volume within the designed vehicle tankage. Mission requirements dictate which of these parameters is most important. Propellant choice is often a matter of compromise between the desire for high performance and the need for ease in propellant handling and low cost. A summary of propellant properties, and some of their vehicle applications, is shown in Table 1.
Propellant Performance Drivers. Thrust is produced by using the energy released in the combustion process to accelerate the resulting gases as they flow through the thrust chamber, thereby increasing their momentum. The gross thrust Fg produced by a rocket engine is the difference in momentum between the propellants entering the combustion chamber and the products of combustion exiting the nozzle. Because the mass flow rate wp is constant through the thrust chamber, the gross thrust is directly proportional to the difference in propellant velocities between the chamber inlet Vn and the nozzle exit Ve:
The theoretical maximum exit velocity Ve can be calculated from the following equation:
Ve = exit velocity
k = ratio of specific heats, Cp/Cv
Tc = combustion temperature
Pc = chamber total pressure
Ru = universal gas constant
M = molecular weight
Pe = exit pressure
All properties used in this equation are those of the combustion products. For expansion to vacuum conditions (Pe = 0), the equation simplifies to
When operating in the atmosphere, the net thrust Fn produced by the engine is reduced by atmospheric pressure acting on the external surface of the engine. Thrust is reduced because the cumulative pressure acting on the outer surface of the nozzle acts in a direction opposite to the direction of gross thrust. Thrust is further affected by the nozzle exit pressure through its influence on exit velocity, as seen in Equation 2. Therefore, net thrust is calculated as
Table 1. Propellant Property References
|Oxidizers||Chemical||Molecular||Tboib K||^freeze?||P Va
* vapor? a
|Liquid oxygen||o2||32.00||90.0||54.4||5,200 at 88.7 K||Cryogenic||Widely used for sea level or|
|upper stage boost|
|Hydrogen peroxide||H202||34.016||419||267.4||345 at 298 K||Earth-storable||In-orbit (extended storage)|
|Fluorine||F2||38.00||85.02||53.54||6,500 at 66.5 K||Cryogenic||In-orbit; applications limited|
|due to toxicity|
|Nitrogen tetroxide||N204||92.016||294.3||261.95||95,800 at 293 K||Earth-storable||Sea level boost (Titan IV: LR87)|
|Upper Stage (Titan IV: LR91)|
|In-orbit operations (Shuttle:|
|Fuels||Chemical||Molecular||Tboii, K||-^freeze?||P Va
1 vapor, a
|Liquid hydrogen||H2||2.016||20.4||14.0||202,600 at 23 K||Cryogenic||Boost (Shuttle: SSME) upper|
|stage (Centaur: RL10)|
|RP-1 (kerosene)||CH1.97||175||460-540||225||2,275 at 344 K||Earth-storable||Boost, primarily sea level|
|operation (Delta: RS-27,|
|Saturn V: F-l)|
|Hydrazine||N2H4||32.05||386.66||274.69||19,300 at 344 K||Earth-storable||In-orbit operations (Apollo|
|Service Propulsion System)|
|Monomethyl||CH3NH2NH||46.072||360.6||220.7||60,657 at 344 K||Earth-storable||In-orbit (Shuttle: OMS)|
|Unsymmetrical||(CH3)2||60.10||336||216||1.213 x 106 at||Earth-storable||In-orbit boost (Ariane 4:|
|dimethyl hydrazine||N-NH2||344 K||Viking V)|
|Methane||CH4||16.03||111.6||90.5||33,000 at 100 K||Cryogenic||Boost: sea level or upper stage|
|Propane||CsHg||36.58||231.0||83.4||896,300 at 298 K||Cryogenic||Boost: sea level or upper stage|
|92.5% ethyl alcohol||C2H5OH||41.25||351||150||89,600 at 344 K||Earth-storable||In orbit|
where Ae is the cross-sectional area of the nozzle exit, Pe is the static pressure of the combustion gases at the nozzle exit, and Pa is the atmospheric pressure outside the engine.
Specific impulse, defined as the net thrust produced per unit of mass flow rate, is calculated from
At high altitudes and in space, the atmospheric pressure term goes to zero, and the specific impulse becomes the ”vacuum impulse” (Ivac).
The unburned propellants enter the combustor in a relatively cold, dense state, so the inlet velocity is very low relative to the exit velocity, which is at high supersonic speed. Considering this difference, the relative features that affect propulsion performance, or Isp, can be examined by considering the factors governing exit velocity. Examining Equation 2, the two factors that can vary significantly and have the most influence on exit velocity are total combustion temperature and molecular weight. These are primarily functions of the fuel and oxidizer chemistry and the mixture ratio. Combustion temperature only weakly depends on pressure. As shown in Equation 2, exit velocity (and Isp) increases with higher combustion temperature and lower molecular weight. For a given ratio of fuel and oxidizer, there will be a resulting set of equilibrium products at a defined temperature. Equilibrium conditions are usually satisfied in conventional liquid-fueled rocket engines. Typical propellants in use today are composed of various combinations of hydrogen, carbon, oxygen, and nitrogen. The products of various propellant combinations are typically composed of mixtures of hydrogen, water, carbon dioxide, nitrogen, and oxides of nitrogen. Hydrogen and water vapor are the lightest products and produce the best performance. Therefore, fuels that are very high in hydrogen content produce better performance when burned with oxygen than hydrocarbons (i.e., kerosene) or hydrazines. Equally important, such fuel-oxidizer combinations typically yield higher combustion temperatures which also results in higher specific impulse. Cryogens. Liquid hydrogen is the highest performing fuel commonly used today. Paired with liquid oxygen, it can achieve specific impulses greater than 470 seconds in vacuum conditions. Liquid hydrogen can achieve such high performance due to the combination of the low molecular weight of its combustion products and high flame temperature, particularly when burned at a near optimum mixture ratio.
Maximum combustion temperature is achieved when the propellants are combined in stoichiometric proportions. For hydrogen and oxygen, this translates to a ratio of 8:1. For maximum performance, however, oxygen and hydrogen are not burned in stoichiometric proportions because that would excessively increase the molecular weight of the exhaust gases. It would also result in a chamber temperature so high that many serious hardware difficulties would be faced. For most practical applications, hydrogen and oxygen are combined in a ratio between 5 and 6:1, the range that provides the optimum combination of low molecular weight and high but acceptable combustion temperature.
Liquid hydrogen is used in applications that require high performance, such as the Space Shuttle and the Centaur upper stage, but it does possess disadvantages. The density of liquid hydrogen is extremely low, about 4.4 lb/ft3 at saturation conditions. This value may be compared with that of another common fuel, kerosene, which has a density of about 50 lb/ft3 (7). The bulk density of a liquid oxygen-kerosene fueled system operating at a typical ratio of 2.72:1 is about 2.8 times that of a comparable liquid oxygen-liquid hydrogen fueled stage operating at a ratio of 6:1. The specific impulse achievable with kerosene is about 25% less than that for hydrogen at typical launch conditions. This results in a propellant mass requirement that is 25% less for hydrogen but is not sufficient to compensate for the large difference in bulk density between the two fuels. A liquid hydrogen rocket, therefore, requires much more tankage volume than a comparable kerosene-fueled system. For a first-stage propulsion system, the large size of hydrogen tanks may make the overall diameter of the stage too large, creating excessive drag losses, or may present significant structural challenges. For this reason, many boosters such as the Saturn V, Atlas and Titan noted before, have used hydrogen-fueled upper stages, for which the propellant load is comparatively small, and noncryogenic first stages.
Two other cryogenic fuels that have been studied for rocket propulsion applications are methane (CH4) and propane (C3H8), both hydrocarbon compounds. Although both propellants are classified as cryogenic, the boiling point of propane is nearly high enough to permit that fuel to be space-storable. It is gaseous at room temperature and ambient pressure but liquid in the colder conditions possible in space. LOX/CH4-fueled engines, in particular, have received attention as possible propulsion systems for Mars missions due to the possibility of using hydrogen and Martian carbon dioxide to form methane. LOX/CH4 engines operating at a mixture ratio of 3.5:1 produce an Isp of about 390 seconds under vacuum conditions; LOX/C3H8 engines, by comparison, yield a vacuum Isp = 386 s with a mixture ratio of 3.2:1 (8). No propane- or methane-powered engines have yet flown on a launch vehicle or spacecraft.
Kerosene (RP-1). Kerosene fuel, otherwise known as RP-1, affords lower performance than liquid hydrogen but offers greatly increased bulk density and ease of handling. Kerosene is Earth-storable, which means that it is in the liquid state under ambient conditions at sea level. Unfortunately, the average molecular weight of the combustion products of kerosene is much greater than that of hydrogen which makes the optimum specific impulse available from a kerosene rocket relatively low. A kerosene/LOX rocket produces an optimum vacuum Isp of around 380 seconds, whereas for a LOX/LH2 rocket the figure is 474 seconds (8). These Isp figures and all quoted below are for the following conditions: Pc = 1000 psia, Pa = 0 psia, AR = 150, where Pc is chamber pressure, Pa is atmospheric pressure, and AR is the nozzle area ratio (Aexit/Athroat). Kerosene is best used when high thrust and ease of propellant handling, rather than performance, are the most crucial design factors.
Alternative Oxidizers. Hydrogen peroxide (H2O2) has been used as an oxidizer since the early days of rocketry. Decomposed hydrogen peroxide powered the turbopumps on the German V-2 (9). NASA has even used hydrogen peroxide in an auxiliary rocket engine for training. The NASA NF-104, used for training X-15 rocket plane pilots, was powered by a Rocketdyne AR2 engine that burned hydrogen peroxide and jet fuel. This engine used a catalyst to decompose the peroxide that was then used to power the turbine and was discharged overboard. The engine was used for several flights until the X-15 program was discontinued.
Although liquid oxygen is by far the most common oxidizer for rocket engines, there are many other choices. Experimentation with cryogenic fluorine (F2) as an oxidizer was carried out in the 1960s due to fluorine’s extremely high performance characteristics. For expansion to a vacuum, an LF2/LH2 rocket can produce an Isp nearly 20 seconds higher than a LOX/LH2 rocket (8). Fluorine has a major drawback for use as a rocket fuel, however, and that is its toxicity. One of the products of combustion is hydrofluoric acid, a poisonous substance. Primarily because of this threat of pollution, fluorine rockets have never advanced past the developmental stage.
Storable Propellants. Uses for which propellants need to be stored on board the vehicle for long periods of time or when launch preparation times are necessarily short are good candidates for storable propellants. One early application of storable propellants arose on the Titan II vehicle. Engine development for this vehicle began in 1960. Originally developed as an ICBM, the Titan II could be launched in about 1 minute, compared to the 15 minutes required to launch its LOX/RP-1 powered Titan I predecessor. The Titan II was fueled by nitrogen tetroxide and Aerozine-50, propellants which were hypergolic as well as storable (10). The current configuration of the first stage engine, the LR87, is flown on the Titan IV vehicle. It develops 548,000 lb of vacuum thrust with a vacuum Isp of 302 seconds (5).
Hypergols are common among storable propellants. In many cases, the engines must be ignited in space, must be operated numerous times, or used under other conditions that make highly reliable operation necessary. Elimination of the igniter removes one significant source of uncertainty from the hardware. The OMS engines and reaction control system for the Space Shuttle, as noted earlier, operate on nitrogen tetroxide and monomethyl hydrazine (MMH). This propel-lant combination has a higher bulk density than either LOX/LH2 or LOX/RP-1, although its performance is modest. N2O4/MMH engines operating at a ratio of 2.37:1 generate an optimal vacuum Isp of 341.5 seconds, lower than hydrogen- or kerosene-powered engines. Another combination of storable propellants uses nitrogen tetroxide as the oxidizer and a 50/50 mix of hydrazine and unsymmet-rical dimethylhydrazine (UDMH). This combination, provides an Isp less than 1 second greater than N2O4/MMH engines and has nearly the same bulk density (7). Most small-thrust, in-orbit applications, in which low performance is not a significant detriment and where it may be necessary to store propellants onboard the vehicle for an extended period of time, commonly use storable propellants.
Optimization of the propulsion system implies minimizing the total amount of propellants that must be carried on the vehicle to satisfy the required thrust in all conditions across the vehicle trajectory. This typically is achieved by maximizing Isp across the flight trajectory, which requires considering several design factors. The following discussion will consider the effects of chamber pressure and the nozzle expansion ratio on the engine performance for applications ranging from sea-level launches to space operations in vacuum.
The net change in propellant velocity is a direct measure of the thrust produced per unit mass burned. Several factors influence the performance that can be achieved with an engine in a specific application. These factors include the selected propellant combination, the combustion chamber pressure, and the exit pressure to which the combustion gases are expanded. The propellant combination and the pressure at which they are burned determine the temperature and properties of the combustion products. The nozzle exit pressure is constrained either by the ambient atmospheric environment or by physical limitations of the spacecraft envelope. For launch vehicles, atmospheric pressure is the constraint, whereas at high altitude or in space, the vehicle envelope limits the allowable nozzle geometry.
The nozzle exit velocity Ve in Equation 2 is a function of nozzle exit pressure. For a constant propellant flowrate and a fixed exit area, as chamber pressure is increased, the throat area needed to pass the mass flow decreases and the nozzle area ratio increases, resulting in higher exit velocity. Therefore, the net effect of increasing the chamber pressure in an engine designed for a specified limiting envelope is an increase in specific impulse that yields a reduction in the fuel required to achieve a given level of thrust.
For application in booster systems, matching exit static pressure with local ambient pressure is typically the controlling restraint on nozzle design. It is most severe at launch and rapidly decreases with altitude. In practice, nozzles are designed for more than ideal expansion at launch (overexpansion) because this yields higher exit velocity and some gain in performance integrated over the flight trajectory. This is a useful design approach because it also reduces the amount of underexpansion that occurs at higher altitude. Care must be taken when designing overexpanded nozzles because too low an exit pressure results in flow separation within the nozzle. Separated flow is very detrimental because the flow is typically unstable and results in uncontrolled variations in the thrust direction. It can also have a significant negative effect on thrust.
Figure 1 shows the trade-off between area ratio and gross thrust (i.e., effects of external pressure not shown) for a LOX/hydrogen engine operating at sea level at a mixture ratio (O/F) = 6:1, where combustion and nozzle performance are ideal and chamber pressure varies from 5-20 MPa. This figure, generated by NASA’s Chemical Equilibrium Applications Program (11), shows the variation in impulse as a function of area ratio at the area ratio at which the nozzle exit pressure is 0.4 bar (note: 1 MPa = 145.03 psia; 1 MPa = 10 bar). This exit pressure is approximately the lowest pressure at which an axisymmetric bell nozzle can be operated at sea level without risk of internal flow separation. As shown in this figure, the final exit area ratio increases from 14 to 40 as the chamber pressure is increased from 5 to 20 MPa. The corresponding specific impulse increases from 397 to 433 seconds, indicating the benefit of initially increasing engine chamber pressure for sea-level launch vehicles.
Engine impulse efficiency is also an important parameter in defining performance. It is the product of the main chamber combustion efficiency and the primary nozzle expansion efficiency. Combustion efficiency measures how thoroughly the propellants mix and burn. Less than 100% efficiency means that some of the fuel or oxidizer leaves the engine unburned, resulting in a loss of thrust. For oxygen and hydrogen propellants, the combustion efficiency for a well-designed system is generally 98 to 99%. For oxygen and kerosene propellants, the practical limit for combustion efficiency is generally 95 to 96%. Nozzle expansion efficiency reflects losses in thrust due to less than ideal expansion of gases in the nozzle. The primary reasons for the loss are gas friction on the nozzle surface and divergence of the flow at the nozzle exit; this means that not all of the flow is aligned parallel to the direction of thrust. Typical nozzle expansion efficiencies are generally from 98 to 99%, depending on several factors, including operating nozzle pressure ratio and the design area ratio.
Figure 1. Variation of impulse with area ratio and chamber pressure.
Another way of looking at trade-off factors for engine performance is to examine the chamber pressure and specific impulse for expansion to a fixed exit pressure. First-order performance trends, as a function of chamber pressure and nozzle exit pressure, are shown in Fig. 2a, for hydrogen and Fig. 2b, for kerosene fuels. The theoretical specific impulse information in these figures was also generated by the NASA Chemical Equilibrium Program. This theoretical performance was adjusted according to the indicated constant values of combustion efficiency and nozzle expansion efficiency and further corrected to account for the external pressure on the engine. Performance is shown for three nozzle discharge pressures Pexit. The Pexit = 1.0 bar line represents the nozzle expansion ratios required to yield a sea-level atmospheric exhaust pressure. This is the exit condition generally required to maximize sea-level thrust. This is desirable and typical for booster applications. The Pexit = 0.3 bar line represents the maximum sea-level expansion ratio that can be sustained without nozzle separation. This is desirable and typical for engines that must start at sea level but also operate at high altitudes. The Pexit = 0.1 bar line is representative of upper stage expansion ratios that balance performance, weight, and engine size. The specific impulse curves in Fig. 2a show that Isp at sea level is reduced if the nozzle expands to a pressure below sea-level atmospheric pressure. This is due to the “negative” thrust effect of the pressure difference (internal minus external pressure) acting on the nozzle. This effect is predicted by Equation 5. In vacuum conditions, the effect of expanding to lower exit pressure is to produce higher performance due to the higher exit velocity produced. In this condition, the external pressure effect has disappeared, and the thrust due to propellant acceleration governs performance.
Figure 2. (a) Hydrogen/oxygen performance trends (b) Kerosene/oxygen performance trends.
These performance estimates are presented only for initial screening. The secondary effects of changes in combustion efficiency and nozzle efficiency that depend on each individual design are important and should be investigated and optimized for each individual application.
It is apparent that increasing chamber pressure is beneficial for both launch vehicle applications and for space applications (limited by nozzle envelope constraints). So why not just increase combustion chamber pressures to reap the apparent benefits? Several factors weigh against this, including increased chamber cooling requirements, increased propellant pumping power, and increased chamber structural weight.
The objective in this section is to provide an overview of the thermodynamics that influences the configuration and optimization of liquid-fueled rocket cycles. The intent is to provide insight into the fundamental differences and inherent capabilities of different cycle approaches. Pressure-fed engines consist ofa single cycle class. These are typically much simpler than pump-fed systems. Some of the issues specific to pressure-fed engines are covered in the section on Propel-lant Supply Systems.
Cycle Types and Configurations. Pump-fed liquid rocket cycles are defined by two configurational variables. The first cycle configurational variable is the energy source for the turbine drive. The turbine energy source can be from an auxiliary combustion device such as a preburner or a gas generator or from the main combustion chamber, either directly by extracting combusted propellants or indirectly by heat transfer through the chamber walls. The second cycle con-figurational variable is the turbine discharge location. Historically, there are two options for turbine discharge flow. If the turbine discharge is to a high-pressure region, specifically the main combustion chamber, the cycle is referred to as a “closed” cycle. If the turbine discharge is to a low-pressure region, generally overboard or into the nozzle skirt, the cycle is referred to as an “open” cycle.
Figure 3 is a summary of eight possible configurational options and includes the common names of each cycle. Also included are the options for turbine working gas supply, propellant limitations, and examples of operational engines of each cycle type. Five of the cycles have been developed into operational engines. This section examines the three most common cycle options. For simplicity, the supporting engine schematics do not include propellant boost pumps and are examined with separate turbopumps for the fuel and oxidizer. The schematics include the minimum valve complement required for engine start-up and control.
Figure 3. Turbopump power options for pump-fed rocket engines.
The basic propellant supply approach for a pump-fed rocket is illustrated in Fig. 4. The propellants are increased in pressure using single- or multiple-stage pumps. A single- or multiple-stage turbine supplies the pump power. Because the propellants for an open cycle are pressurized only slightly above chamber pressure, pump work is minimized. A turbine pressure ratio of five or greater is possible because of the low-pressure exhaust. For a closed cycle, the turbine drive flow is discharged into the main chamber, which is at a relatively high pressure. This generally limits the turbine pressure ratio to two or less to avoid excessive pump discharge pressures. For either the open or closed cycle approach, it is necessary to introduce energy into the turbine working fluid before expansion through the turbine. Depending on the option selected to provide this turbine energy, the cycle definition is different. The three common thermodynamic cycles for liquid rocket engines are expander, gas generator, and staged combustion.
The expander cycle, Fig. 5a, is a cycle in which hydrogen, or some other fuel, is used to cool the thrust chamber and nozzle regeneratively. Thermal energy absorbed during cooling of the chamber and nozzle heats the hydrogen fuel. The heated hydrogen, now in a gaseous state, passes through turbines and powers the pumps. In this engine cycle, the turbine gases are routed to the injector and main chamber, where they are combusted and expanded through the nozzle. The thrust chamber and nozzle heat transfer limits the energy available for the expander cycle. This limits chamber pressure potential to about 10 MPa (1500 psia).
Its simplicity is the major benefit of this cycle. There is no subsystem needed to provide the energy to drive the turbines. By the same token, its main drawback is that the energy of the turbine working fluid is limited and leads to relatively low chamber pressure.
Figure 4. Pump-fed propellant supply schematic.
The gas generator cycle, Fig. 5b, is an open cycle configured so that a minimum fraction of the propellants is delivered to the gas generator combustion device before being directed to the high expansion ratio turbine. The pressure of the turbine discharge is less than the main combustion chamber pressure, and therefore this flow must bypass the main combustion chamber. The gas generator cycle dumps the gases used to power the turbopumps overboard or into the divergent section of the nozzle. The chemical energy released during combustion in the gas generator is restricted by the temperature limit of the turbine. Chamber pressure for a gas generator cycle is selected to optimize total engine performance, which includes both the higher performance main engine flow and the lower performance turbine discharge flow. This performance optimum generally occurs at 10 to 17MPa chamber pressure, depending on propellant selection, and the overboard flow is generally less than 4% of the total engine flow.
Figure 5. Schematics of liquid rocket engine cycles: (a) Expander cycle. (b) Gas generator cycle. (c) Staged combustion or preburner cycle.
The propellants for the gas generator are typically the same as those burned in the main chamber, though decomposition of hydrogen peroxide has sometimes been used, most notably in the German V-2 (9). The Rocketdyne RS-68 gas generator engine recoups some of its performance loss by using liquid hydrogen as a fuel, but its 410-second average vacuum specific impulse is still much lower than that of a comparable staged combustion engine.
The use of gas generators allows increasing the chamber pressure above that which is possible in the expander cycle. The cost of this is added complexity and some loss of thrust from dumping turbine discharge gases.
The staged combustion cycle, Fig. 5c, is a closed cycle configuration such that portions of the propellants are burned fuel-rich in preburner combustion devices upstream of the turbines. This heated mixture of fuel and combustion products is expanded through the turbine and fed into the main combustion chamber. In the SSME, the primary operating example of this cycle, approximately 80% of the fuel flows through the fuel turbine. The system must be balanced between the desire for high chamber pressure and the need to limit turbine inlet temperature to an acceptable value dictated by hardware requirements. Turbine inlet temperature is controlled by the amount of oxidizer that is fed into the preburner. For the SSME example, the mixture ratio in the pre-burners is of the order of unity. The performance of the staged combustion cycle begins to become hardware-limited between 20 and 24 MPa (3000 and 3500 psia) chamber pressure.
Newer designs would probably achieve pressures near 27 MPa. Use of full flow, ultrahigh performance designs, where a low mixture ratio, fuel-rich flow drives the hydrogen pump, and a high mixture ratio, oxygen-rich flow drives the oxidizer pump, can achieve thrust chamber combustion pressures in the 27 to 38 MPa (4000 to 5500 psia) range. Use of cooled preburner and turbine hardware (including turbine blades) can extend the combustion pressure to
Table 2 illustrates some of the basic cycle parameters for a variety of liquid rocket engines in current use, including thrust chamber operating conditions, type of cycle, and propellant pump conditions.
Propellant Supply Systems
The propellant supply system consists of the various components that store the propellants in the vehicle and deliver them in a controlled manner to the engines. These components include the propellant tanks, valves, pumps, plumbing, and pressurization subsystem. Depending on the type of propellant pressurization (i.e., pumped or pressure-fed), the propellants are supplied as liquids from the vehicle tankage to the inlets of the main pumps or to valves that control the flow into the engine.
Table 2. General Cycle Characteristics of Some Current Rocket Engines
|Engine||Vehicle||Propellants||Cycle||Pc, psia||Tc, °R||Pump, hp||Pump speed, rpm|
|SSME (Blk II)||Shuttle||LOX/H2||Staged combustion||3000||6500||69,000 fuel, 25,150 LOX||35,000 fuel, 24,000 LOX|
|RL10A-4 (Pratt||Atlas-Centaur||LOX/H2||Expander||565||6100||740 fuel,||35,500 fuel,|
|Whitney)||160 LOX||14,200 LOX|
|Vulcain (SEP)||Ariane 5||LOX/H2||Gas generator||1600||6300||16,100 fuel, 5,100 LOX||33,500 fuel, 13,800 LOX|
|RD170 (NPO||Zenit||LOX/RP-1||Staged combustion||3500||6900||55,000 fuel||17,000 both|
|Energomash)||(Main + kick), 123,00 LOX||(single shaft)|
|RS-27A||Delta II,||LOX/RP-1||Gas generator||700||6400||2000 fuel,||6800 fuel,|
|(Rocketdyne)||Delta III||1100 LOX||6800 LOX|
|LR-87-AJ-11||Titan IV||^CVAerozine||Gas generator||860||5900||3800 fuel,||9500 fuel,|
|(Aerojet)||50||4000 LOX||8700 LOX|
The parameter conventionally used to quantify the inlet condition is the net positive suction head (NPSH), the total head difference (i.e., pressure plus velocity head) between the propellants at the inlet and their corresponding vapor pressure. It is desirable that the NPSH at the pump inlet be sufficient to avoid cavitation in the pump. One way to achieve this is by maintaining high propel-lant tank pressure. This approach may result in excessively heavy tanks, particularly for propellants with high vapor pressures. In addition, the empty tank volume, called “ullage,” must be at the same pressure as the propellants in the tank. As liquid is withdrawn from the tanks, pressurized gas must be injected to replace the liquid volume. This ullage gas can be either heated and vaporized propellant or an inert pressurized gas. At higher tank pressures, this represents a significant mass of pressurizing gas, particularly as the propellant level in the tank approaches “empty.”
One design option is to supply the propellants at pressures that are a compromise between minimizing the tank ullage pressures and maximizing the net positive suction head. The vehicle prefers low ullage pressure so that the tank walls are thinner and therefore lighter in weight. This creates a lighter or smaller vehicle for the same propellant load and translates into larger pay-loads delivered to orbit. Alternatively, for a given engine operating chamber pressure, pump-fed engines need high enough inlet pressures to prevent cavi-tation in the pump. Selecting a propellant with a vapor pressure too low or too high can severely affect the overall tank and propellant feed supply design. It could require higher ullage pressures that lead to heavier tanks on the vehicle.
A second option is to design the tanks for low ullage pressures to minimize tank weight and to satisfy the required NPSH level by adding propellant boost pumps to increase the net inlet pressure to the main pumps. This adds complication to the system but can provide significant overall benefit to the vehicle. Boost pumps are typically small axial flow pumps designed to operate at low pressure, referred to as “inducers”; they are placed either at the tank or at the inlet of the main pump. The inducer has the effect of increasing the net positive suction head at the main impeller inlet by boosting the total fluid pressure.
Pressure-fed rocket engines can be designed to supply a single propellant (i.e., hydrazine thrusters) or as a bipropellant system, like the space shuttle OMS. Tank pressurization is achieved via complementary pressurization tanks or bottles of an inert gas such as helium or nitrogen. Regulator valves are used to control the gas pressure relative to the fluid pressure in the main tanks. The system must be designed to compensate for the pressure losses between the pressurization tanks, valves, main tanks, propellant lines, and the main chamber injector. Pressure-fed propellant supply systems are normally limited to engine operating pressures of less than 3 MPa.
The propellants are supplied to the engine inlet valves, control valves, or pump inlets via fixed or flexible propellant feed lines. In some engine designs where vernier or auxiliary engines provide steering the engine does not move. In that case, the engine propellant supply lines are usually hard-mounted to the inlet valves or pumps. When the engine is vectored or gimbaled to provide flight control, the propellant supply will use a flexible duct line that permits “jointed” movement in two planes. In some cases (e.g., Russian RD-170 LOX/kerosene engine), the propellant supply lines are fixed, and the thrust chamber gimbals to provide vectored flight control.
Liquid Propellants Turbopumps
The function of the rocket engine turbopump is to receive the liquid propellants from the vehicle tanks at low pressure and supply them to the combustion chamber at the required flow rate and injection pressure. The energy to power the turbine is provided by the expansion of high-pressure gases that are often mixtures of the propellants being pumped. This section relies heavily on Reference 12.
The liquid rocket engine turbopump is a unique piece of rotating machinery. The turbopump typically pumps cryogenic liquids and is driven by high-temperature gases, creating large temperature differentials between the pump and turbine. The pump must avoid cavitation while pumping relatively high-density fluids at low inlet pressures and deliver them to the thrust chamber at very high pressures across a relatively wide throttling range. The turbine is often driven by fuel-rich combustion products that have very high available energy and heat capacity levels. The turbopump is optimized for performance and weight within the minimum possible envelope size to facilitate engine packaging. The bearings normally operate in the environment of the propellants being pumped, which have minimal lubrication characteristics. The static and dynamic seals must preclude mixing propellants within the turbopump, which would result in burning and catastrophic failure.
Engine Requirements. The type of engine cycle has the most significant influence on the turbopump requirements and configuration. Other major engine factors that significantly influence the turbopump configuration are the types of propellants, the propellant inlet conditions, and the engine throttling requirements. Variations in density produce significantly different pump head rise requirements and large differences in volumetric flow, that is, low-density propellants require a much higher head rise to develop the same discharge pressure (head rise = pressure rise/density, AH=AP/p). The variations in the combined propellant available energy also have a significant influence on turbine design.
The pump suction performance requirement is its ability to operate at the available NPSH without cavitation sufficient to affect its ability to develop the required discharge pressure and flow rate.
The engine throttling requirements define the range of flow and discharge pressure that the turbopump must deliver in stable operation. The engine start and shutdown characteristics must also be considered to prevent unstable tur-bopump operation caused by cavitation or stall.
When the engine requirements are established, the turbopump configuration is selected based on optimizing the pumps for each propellant, the turbine for the drive gas available energy, and the mechanical design arrangement for life, weight, and producibility considerations. Maximum pump speed is generally limited by the suction performance requirements to avoid cavitation. The optimum turbine speed for maximum efficiency and minimum weight is generally higher than the high-density fluid pump speed. Maximum turbine efficiency requires a certain pitch-line velocity, which is the product of the shaft speed and the turbine diameter. The minimum weight turbine has the highest speed and smallest diameter within the structural and mechanical arrangement limitations.
Earlier engines with small power requirements sometimes employed a gearbox to match the speeds of the pumps and turbine better, but at the very large power levels of launch systems, pumps are typically direct driven by the turbine. Therefore the turbine must satisfy the power requirements of the pump at the same shaft speed.
Pumps. Inlet conditions (NPSH), discharge pressure, flow rate, and operating range must all be satisfied by the pump configuration. A parametric analysis is performed to select the best speed, diameter, and number of stages compatible with the turbine and mechanical design considerations (Fig. 6).
The pump inlet diameter selected is generally based on the available NPSH. Test experience has been accumulated on inducers to correlate their suction performance as functions of the NPSH, the fluid inlet meridional velocity Cm, and the inducer flow coefficient j. The inducer diameter (inlet area) is selected to limit the fluid meridional velocity so that the available NPSH/2gcCm is equal to or greater than three velocity heads for water, two for LOX and one for LH2. The variation in the empirical limit accounts for the difference in the thermodynamic suppression head among water, LOX and LH2. The limit is also a function of the inducer flow coefficient, which is defined as the meridional velocity divided by the inducer tip speed, j = Cm/Ut.
Figure 6. Performance and structural limits establish basic design.
When the inlet diameter is selected, the shaft speed is selected to limit the inducer tip speed to approximately 550 ft/s in LOX and 1100 ft/s in LH2. The tip speed limit is for controlling the tip vortex cavitation energy. The blade thickness must also increase with increased tip speeds to react to the centrifugal and pressure loading. This results in reducing the flow passage area and, therefore, lowers the suction performance. The pump suction specific speed is expressed as Ss = AVQ/(NPSH)3/4, which is a measure of the pump’s ability to operate at low inlet NPSH without cavitation sufficient to cause head loss. A 50% NPSH margin is generally selected during the design process for long-life rocket engine applications. Cavitation, in addition to decreasing the pump discharge pressure and efficiency resulting from the formation of vapor bubbles, can cause significant structural damage when the vapor bubbles collapse (implode), particularly in high-density fluids. Inducer technology development has been a key state-of-the-art advancement for increasing pump speed, decreasing turbopump weight, and increasing safe operating life (Fig. 7).
Figure 7. Rocketdyne suction specific speed history.
Required pump head, which is a function of the required discharge pressure, the available inlet pressure, and the propellant density [AH = (Pd — Pin)/p], is the major factor in selecting the pump configuration. The head coefficient is a function of the pump type and establishes the required pumping element diameter and number of stages for a given shaft speed. The main pumping element may be a centrifugal, mixed, or axial flow type (Fig. 8).
For the combination of very high flow rate and large head rise usually encountered in moderate and large engine applications, centrifugal pumps are typically chosen as the most appropriate type of pump. In this pump type, pressure rise is achieved by a combination ofacceleration and centrifugal force, as the fluid flows along a curved flow path. The pressure rise achieved depends on both the characteristics of the pump (i.e., impeller design and rotational speed) and on the density of the fluid. In general, pressure rise in a single centrifugal stage is proportional to the product of the fluid density and the square of the tangential velocity at the outer rim of the impeller (tip speed). Thus, the higher the density, the lower the required tip speed to achieve a given pressure rise. In general, the tip speed and impeller diameter are well within aluminum and nickel-base alloy structural limits. The head requirements for low-density fluids, such as LH2, are very high and typically require several stages to develop. An axial flow main pumping element was selected for the J-2 LH2 pump because of its intermediate specific speed and the narrow throttling range requirements. The 200,000-ft head requirement for the SSME high-pressure fuel turbopump dictated a three-stage centrifugal pump with impellers operating at 2,000 ft/s tip speed. Titanium, which has a higher strength-to-weight ratio than high-strength nickel-base alloys, was required by the high tip speed.
Figure 8. Pump types (photos).
The pumping power requirement depends on the flow rate, pressure ratio, and thermal efficiency and can be calculated from
where wp is the propellant flow rate,Hi is the inlet enthalpy,Hoactual is the actual exit enthalpy, Hoideal is the ideal isentropic exit enthalpy, and z is the pump efficiency. The efficiency is determined by the internal aerodynamics of the pump and is typically characterized for a defined pump geometry as a function of speed and pressure ratio or speed and flow rate. Pumps can be designed to operate at high efficiency at their design conditions, but efficiencies tend to be lower if the pump is operated at conditions far from design. It is desirable to design the pump to match best the cycle conditions over the engine operating envelope, maximizing the efficiency to minimize power requirements.
The Space Shuttle Main Engine (SSME) represents the extremes of current turbopump requirements. The design requirements of the liquid hydrogen fuel pump at maximum power are 162lb/s flow rate at a discharge pressure of 6400 psia. The liquid oxygen pump delivers 1161 lb/s at 7300 psia at maximum power. It is noteworthy, on examining the operating characteristics shown in Table 2, that hydrogen fuel represents about 12% of the total propellant flow but requires about 74% of the total pumping power. This is the result of the relatively low density of liquid hydrogen. The table also shows that the speed of the hydrogen pump is 50% greater than that of the oxygen pump, again because of the very large head rise necessary to produce the desired pressure rise in a low-density fluid.
An ideal fluid with regard to pumping requirements would be inert with respect to the pump materials, would have a relatively high density, and would be a good lubricant and coolant for use in the pump bearings. To illustrate the contrast in propellant characteristics, consider some of the propellants in current use (Table 3).
Optimizing the pump efficiency, which is a measure of the work out/work in, can also influence the shaft speed and specific speed selected. Figure 9 shows the relationship between efficiency and specific speed. Small flow rate pumps are generally less efficient than large flow rate pumps because the clearance and surface-finish-related losses cannot be scaled with size.
Turbines. The turbine must supply the required power to drive the pump, using the drive gas provided by the selected engine cycle. It is desirable that the working fluid be at the maximum possible temperature to provide the maximum work potential. However, the turbine inlet temperature is limited by the material and structural capabilities of the turbine because it is very desirable to avoid cooling the turbine. Axial flow turbines tolerate high gas temperatures better than radial inflow turbines and generally have much lower thermal stresses. For this reason, turbines for rocket engines are generally the axial flow type and may use either impulse or reaction type aerodynamics. Small engines will usually use impulse turbines because of their ability to use partial admission nozzles and avoid significant leakage losses. Large thrust engines may use reaction turbines, if the working flows are large enough to avoid requirements for small blading. Small reaction turbines are difficult to use because they require extremely tight clearances and seals to avoid significant flow leakage and have large attendant performance losses.
Table 3. Propellant Characteristics of Importance for Pumps
|Inert with regard to most pump material
Ambient temperature storable
Relatively good lubricant
Density ~ 50 lb/ft3 Nearly ideal for ease of pumping
|Causes embrittlement in
nickel alloys because of
chemical reaction Deep cryogen, liquid at 20 K Materials must be cryogenic-
capable (not brittle at
hydrogen temperature) Virtually no lubrication
capability, although very
good coolant Density is very low ~4.4 lb/ft3 Difficult propellant for pump
|Highly reactive with many materials
Cryogenic, liquid at 90 K Materials must be cryogenic-
capable and resistant to
oxidation Poor lubricant qualities high
density ~ 71 lb/ft3
Easier pumping than hydrogen, but requires careful material choices
Figure 9. Variation of pump efficiency with specific speed.
The operating characteristics of turbines are defined by the power produced as a function of the flow rate and the pressure ratio across the turbine. This is determined by the product of the flow rate and the enthalpy change as the working fluid expands in the turbine. This expansion process takes place as flow is accelerated across the nozzles or vanes and converts pressure to kinetic energy, and momentum is removed as forces acting on the rotating blades. Another way of looking at turbine performance is that it depends on three variables: the available energy content of the gas, the blade tangential velocity U, and the number of turbine stages. The available energy for the turbine pressure ratio can be expressed as an ideal velocity C. The turbine velocity ratio U/C is used to characterize these two variables empirically versus the turbine efficiency.
The ideal velocity can be distributed among the turbine stages by either a pressure-compounded or velocity-compounded design. The major difference between these two turbine designs is where the expansion occurs in the stationary blade rows. For the velocity-compounded turbine, all of the expansion occurs in the first stationary blade row, and for a pressure-compounded turbine, the expansion is distributed between the stationary blade rows. For high U/C designs, the turbine efficiency can be further improved by having some of the expansion (reaction) take place in the rotor blades, as illustrated in Fig. 10. The design selection is made to maximize the turbine efficiency and minimize the weight, compatible with the selected shaft speed. In general, when a direct-drive turbopump configuration is selected, the shaft speed is less than optimum for the turbine, and additional stages must be added to use the available energy.
Figure 10. Velocity ratio vs. efficiency for impulse and reaction staging.
The blade tip diameter is selected to optimize the U/C for efficiency within the blade height-to-diameter performance limits and within the tip speed structural limits. If the blade height-to-diameter ratio becomes too small, the tip clearance and secondary flow losses become large and decrease the turbine efficiency. The tip speed structural limit is based on the centrifugal pull that can be carried at the base of the blade airfoil for the selected material. Partial admission turbines are selected when the shaft speed is too slow and the blade height-to-diameter ratio becomes too small to obtain the desired U/C. The blade diameter is enlarged to increase U, and the arc of admission is decreased to maintain the blade height at an acceptable height-to-diameter ratio.
Mechanical Design. The mechanical design is a compromise that involves many contributing factors. Major factors that influence the mechanical design are power transmission, rotor dynamics, axial thrust balance, selection of bearings and dynamic seals, and thermal considerations.
The shaft diameters, splines, and couplings must all be large enough to transfer the torque, which is a function of the speed and horsepower. This establishes the shaft diameters and the minimum allowable diameter of the bearing inner race, depending on the axial stations selected along the shaft. Rotor critical speeds are a function of the rotor mass distribution and stiffness, bearing locations and spring rate, and the housing stiffness. For rolling element bearings, the product of the inner diameter and the shaft speed (DN) is used as a measure of the bearing internal loading. Empirical life limits have been established for DN as a function of the propellant or lubricant used to cool the bearing. Based on these interacting limitations, the bearing locations are selected to keep the operating speed range clear of critical speeds and minimize the bearing bore diameter to maximize bearing life.
From the standpoint of critical speed, inboard bearings decrease the bearing span and increase the first critical speed, so long as the overhang does not exceed approximately one-half the span length. From a bearing standpoint, the most desirable location is outboard, so that the bearing size and geometry can be optimized independent of the required shaft diameter. Bearings outboard of the turbine generally require additional cooling, dynamic seals, and support structure, which add complexity to the turbopump design.
Rotor stability is also a major factor in selecting the bearing locations and types of dynamic seals. Additional support stiffness and damping can be provided by the dynamic seals to raise the critical speeds and stabilize the rotor to prevent subsynchronous whirl. The rotor axial thrust is the other major factor that influences bearing design. The labyrinth seal diameters in the pumps and turbine are selected to minimize the net rotor thrust to which the bearings must react.
The rocket engine turbopump, in addition to being a high energy/weight ratio machine, must be designed to operate with the pump at cryogenic conditions and the turbine at high temperature. This requires design concepts that provide thermal growth flexibility while reacting to large torques and separating loads.
Materials. Aluminum alloys, stainless steels, high-strength steels, nickel-base alloys, cobalt-base alloys, and titanium alloys are all used in the design of rocket engine turbopumps. Complex pressure vessels for applications up to approximately 2000 psi are typically cast of aluminum to use its high strength-to-weight ratio and to avoid welded joints. Nickel-base superalloys, such as Inco 718, are used to cast pressure vessels when higher strength is required. The high strength/weight ratio of titanium is used to obtain the high tip speeds required for LH2 impellers and inducers.
The embrittling effects of gaseous hydrogen limit the materials suitable for turbine components. High-strength superalloys typically must be protected from the environment by copper or gold plating. Turbine blades are directionally solidified and thermally coated to survive heat fluxes 10 times the typical turbojet and blade loads up to 600 hp per blade.
Silver and Kel-F are used in LOX pumps where contact with the inducer or impeller could result in ignition caused by local heat generation. These materials are also used for potential contact with titanium impellers to preclude formation of titanium hydrides caused by heat generation.
Probably the most significant technological advancement to impact future turbopump designs is the development of fluid film bearings. Hydrostatic bearings remove the DN constraints of rolling element bearings, increase the bearing direct stiffness by a factor of 5, and increase damping by a factor of 100. Rotor stiffness can be increased, and the bearings can be positioned to optimize the machinery arrangement for rotor dynamics and performance considerations. A significant technological need for LH2 turbopumps is the development of high strength-to-weight materials. Specific strengths higher than titanium will increase the 2000-ft/s tip speed limit, increase the head generated per stage, and decrease the number of stages required for the desired discharge pressure. Further information on this subject can be found in Reference 13.
Thrust Chamber Design
A thrust chamber is that part of the rocket engine that produces the thrust; it is defined as the section of the flow path enclosure extending from the injector face to some location downstream of the nozzle throat, at which the thermal and structural loads are reduced sufficiently to allow significant reduction in the weight of the flow path structure. The nozzle structure downstream of this location is usually referred to as the nozzle extension. The thrust force is transmitted to the vehicle thrust frame through a thrust structure mounted on top of the injector that allows gimbaling the whole engine. Figure 11 shows a typical thrust chamber configuration found in all flight engine designs today, using an axisymmetric chamber and a bell-shaped nozzle extension. Other configurations such as circular or linear aerospike designs or expansion-deflection nozzle designs have been successfully used in experimental engines but have not matured enough to demonstrate performance benefits in flight hardware compared to the more traditional design.
In the thrust chamber, the propellants are manifolded and injected into the combustor by the injector and are atomized, mixed and burned inside of the combustor. The resulting hot gases are expanded inside the nozzle extension, accelerating the flow to high exhaust velocities and producing the required thrust. In a typical design, about half the thrust is generated by the flow downstream of the thrust chamber throat. The thrust chamber is functionally characterized by components that provide stable and highly efficient propellant combustion and efficient combustor and nozzle cooling. Design activities at the thrust chamber level start out with design trade-offs to define optimum thrust chamber performance and resulting component requirements. Typical system interface issues that must be covered include the injector-combustor interface (”chamber wall compatibility”), trading combustion efficiency and stability versus chamber wall heat flux loading, and the location of the combustion chamber-nozzle extension interface. This usually involves trading off the system weight against performance and manufacturing requirements. For nonhypergolic propellant combinations, such as LOX/RP-1 and LOX/LH2, an igniter is needed to initiate combustion. In most cases, an electrical ignition system is employed. The thrust chamber design must take into account all loads encountered during steady and engine start and stop operation, such as pressure loads, thermal loads, nozzle side loads, as well as taking into account requirements for external loads and geometric interface locations.
Figure 11. Photo of regeneratively cooled thrust chamber (Vulcain).
Figure 11 shows the cross-section in a photograph of a 1600-psia injector/ combustor assembly, which will be used to outline the major design issues and features of a regeneratively cooled high-pressure thrust chamber. The design characteristics shown are similar for all hydrogen/oxygen systems used worldwide, such as the SSME (U.S.), the RD-120 (Russia), the Vulcain (Europe), and the LE-7 (Japan). Most regeneratively cooled chambers are fuel-cooled in a counterflow cooling arrangement; the regenerative section of the nozzle extension is also fuel-cooled. For overall engine performance, it is important that the coolant pressure drop is minimized.
The main design goals and requirements for the injector are high combustion efficiency and stability, providing additional chamber wall cooling, if needed, and sufficient structural strength and life. This leads to the appropriate definition of the geometry (e.g., the manifold shapes), the injector element design, the element pattern, and the materials and manufacturing processes. High combustion efficiency requires a uniform mixture distribution across the injector face and fine propellant atomization. A suitable manifold achieves the uniform mixture distribution and injection pattern design. The selection and optimization of the most suitable element type provides fine propellant atomization. Coaxial injection elements are used for liquid/gas systems like the one shown in Fig. 12 (LOX/GH2) and for gas/gas systems. Liquid/liquid systems such as LOX/RP-1 or MMH/N2O4 usually feature impingement type injection elements, even though coaxial element designs are also successfully used in MMH/N2O4 systems. Unfortunately, many of the injection parameters that provide high performance, such as fine atomization, tend to reduce the combustion stability margin, and stable operation must be achieved by damping the acoustic processes (i.e., resonator cavities in the combustor wall) or by detuning them (i.e., baffles in the injector). Storable propellant combinations like MMH/N2O4 are most susceptible to combustion instability problems, LOX/RP-1 less so. The hydrogen/oxygen propellant combination is the least susceptible to high-frequency instability. To avoid coupling the combustion process with a feed system hydraulic mode (“chugging”), a sufficiently high pressure drop must be designed into the injector element, typically about 20% of the chamber pressure.
Figure 12. Schematic of Vulcain regeneratively cooled thrust chamber.
A typical injector like the one shown in Fig. 11 has about 500 injection elements. The faceplate needs to be actively cooled; this is achieved either by transpiration or regenerative cooling. Most parts of the injector are made from high-strength materials such as Inconel 718 or titanium, and for lower pressure applications, aluminum. The main manufacturing processes used are casting, turning, milling, drilling, brazing, and welding.
The main design goals for the combustion chamber and the nozzle section are nearly the same and may be listed as liner life, structural strength (and life), additional wall cooling requirements, and combustion stability devices (i.e., ”acoustic cavities,” used only in the combustion chamber). This leads to the definition of the hot gas wall contour, the cooling channel design, and the selection of materials and processes. The biggest design challenges are the structural and thermal design of the combustor and the nozzle extension to ensure attaining the required useful life within allowed coolant pressure losses and without exceeding the target weight for the chamber component. In the case ofthe nozzle extension, the superposition of thermal loads and structural loads in the form of side loads, both mainly due to cyclic loading, (engine start/stop) must be covered. If the engine operates on an expander cycle, the heating of the chamber and nozzle coolant must be maximized to drive the turbines at the required power level.
High chamber pressure results in high heat loads (e.g., 100 Btu/in.2-s. at 3000 psia chamber pressure for the SSME) and dictates the use of a high-conductivity copper alloy liner of minimum wall thickness. Today’s high-pressure combustors use the milled channel design, whereas future combustors for expander cycle engines may use a tubular design. After repeated cyclic loading (engine start/stop) and high heat flux/high temperature liner operation, the liner may fail structurally, primarily due to thermal ratcheting which is largely a combination of low-cycle fatigue and thermal creep. Hence, the cooling must be designed to control the maximum temperature levels of the gas-side liner material along the coolant flow path, which allows the required chamber life. For a prescribed hot-gas flow path contour, the cooling channel or cooling tube geometry must be designed to keep the maximum liner temperature below the limit required for life, while not exceeding the available pressure drop. As a typical example, the chamber shown in Fig. 11 has 360 cooling channels and a pressure drop of about 30% of the chamber pressure. Pressure drop increases with chamber pressure, and is about 50% of the chamber pressure for the SSME.
In hydrogen/oxygen systems, a chemical effect called blanching occurs above certain liner temperatures and degrades the liner material’s properties. Similar effects (chemical attack) occur with other propellant combinations, (e.g., MMH/N2O4) whose combustion gases are incompatible with copper alloys and require a coating to allow high-pressure operation. Lower pressures may allow using a nickel liner material that offers a compromise between liner conductivity and susceptibility to chemical attack, depending on the operating regime. In LOX/RP-1 systems, another propellant-related effect occurs below a certain pressure level. During operation, the hot gas wall is coated with a condensed carbon layer, called coating, effectively reducing the heat flux. On the coolant side, propellant material compatibility issues are hydrogen embrittlement of nickel-based alloys or coking of the coolant side wall depending on which pro-pellant is used. These effects lead to increased coolant pressure drop and to liner material temperatures that may result in chamber burn-through.
A milled channel design requires turning and milling operations on the copper alloy (CuAgZr or CuCr), closing the channels, and attaching a structural jacket. Processes used for closing milled channel chambers include nickel elect-roforming optionally reinforced with a welded shell or compression brazing a high-strength jacket directly to the liner. In a different construction approach, tubular chambers are formed by brazing tubes and applying a structural jacket by one of the methods cited, or by using thermal spray techniques to apply an external metal structure. Inlet and outlet manifolds are typically inert gas welded or electron beam (EB) welded to the structural jacket.
Nozzle sections and extensions are basically formed by the same methods as described before. However, stainless steel materials usually have sufficient heat flux capability. Tubular structures are used for lightweight designs, and milled stainless steel liners may be used if weight is of minor importance. Film-cooled and radiation-cooled metallic or ceramic matrix composite nozzle sections may also be incorporated, depending on the application (e.g., the new Rocketdyne RS-68 uses a ceramic ablative nozzle).
Several new systems that could compete with conventional liquid rocket engines for the Earth-to-orbit mission are becoming more practical due to recent advances in technology. First and foremost among these are supersonic combusting ramjets, or scramjets. The concept ofscramjets was developed some time ago, but severe technical obstacles had to be overcome before they could be considered for application. In the 1970s and 1980s, a series of breakthroughs occurred and since then several programs were begun to develop a practical vehicle using this for propulsion. The ability of scramjet engines to obtain their oxygen from the atmosphere leads to a much higher Isp than possible with other liquid systems. For instance, in the 8-10 Mach number range, scramjets have an Isp greater than 3000 seconds (14). This falls off as the Mach number increases. The upper speed limit of the scramjet has not been determined, but theoretically it is above the Mach 20-25 speed range required for orbital velocity. The scramjet benefit compared to a rocket becomes small as its speed approaches orbital velocity, but this is only academic because the flight environment provides nearly insolvable structural and internal and external aerodynamic challenges.
System considerations lead to vehicles using scramjet propulsion that are very much more like airplanes than rockets. This in turn results in their ability to use aerodynamic forces, rather than rocket thrust, to control the vehicles. Besides Isp, which translates to a takeoff gross weight advantage, these vehicles have advantages in flight safety (abort, fly back) and mission flexibility (launch window, orbital offset, and rapid rendezvous). The major disadvantage is technological readiness. Rockets and rocket-powered vehicles were refined and re-refined during the past 40 years. Airbreathing vehicle technology is just now coming of age.
Because the scramjet-powered vehicle obtains all of its oxygen from the inlet, the vehicle integration problem is much more severe than with a conventional rocket vehicle. The forebody of the vehicle is really a part of the inlet, and the aft end of the vehicle functions as part of the nozzle. Among other things, this places operational restrictions on scramjet-powered vehicles, such as Mach number and angle of attack limitations to ensure inlet air capture. Effective scramjet-powered vehicle design requires true synergy, where the engine and vehicle functions blur, and the only true measure of scramjet performance is the overall mission success.
Scramjets cannot operate at low speeds, so engines that are required to operate across a large speed range must include another cycle. Various combinations have been studied, and the rocket-based combined cycle (RBCC) is a popular choice because of its simplicity (i.e., use of a single flow path). RBCC engines use a rocket integrated within the scramjet combustor, as shown in Fig. 13, to provide static thrust for takeoff, subsonic, and low supersonic operation up to ramjet speeds. Following dual-mode scramjet operation to Mach 10-16; the rocket is used first to supplement the scramjet thrust and finally alone to allow orbital insertion and in-space maneuvers. Another approach being considered uses a high-speed turbojet for flight up to Mach 4-5; then the dual-mode scramjet and separate rocket are used for higher speed. This turbine-based combined cycle has additional complexity but uses a more reliable engine cycle (turbojet). The best candidate for reduced cost, improved reliability, and safety for space access is an issue of current studies.
Figure 13. Rocket-based combined cycle.
The scramjet engine consists of an inlet, an isolator, a combustion chamber, a nozzle, and a fuel pump and supply system. These components each have design criteria that define their configurations and performance characteristics. In addition, a scramjet propulsion system is a highly integrated aerodynamic device with strong interactions among components; each component must be designed to satisfy integration criteria for the total system.
The inlet is designed for no “spillage” in the design flight condition because spilled air is an additional drag on the aircraft. This is achieved for two-dimensional mixed compression inlets by configuring the inlet so that the first inlet shock rests on the lip of the opposing lower inlet cowl. For a typical mixed compression inlet, where the forebody of the aircraft acts as a precompression surface, the total compression is roughly split equally between the external forebody and internal to the inlet. The other primary design criterion is that the inlet provides downstream flow conditions to satisfy the requirements of temperature, pressure, and velocity necessary to achieve combustion within the confines of the combustor. Good engine performance also requires relatively uniform air distribution at the combustion inlet. Compression in scramjet inlets must be restricted so that the flow remains supersonic. For example, a Mach 7 scramjet-powered vehicle would have a combustor entrance Mach number of about 2-3. The function of the isolator is to limit the amount of pressure that is fed upstream in the wall boundary layers from the combustor, so that the inlet does not unstart. A series of oblique shocks exist in the isolator as a normal part of achieving the necessary compression.
Downstream of the inlet and isolator is the combustor. Here fuel is injected into the supersonic airflow, mixes and burns within the short time that it takes to traverse the combustor length, typically of the order of milliseconds. The location and amount of fuel injected varies with the operating and flight conditions. At low relative Mach numbers, fuel is supplied from wall injectors near the downstream end of the combustor. As speed increases, the fuel supply will gradually be moved forward until at hypersonic speeds, the fuel will be injected at the downstream end of the isolator. The design of the injector elements is extremely demanding because of the requirement for ultrarapid mixing. Variation of the fuel injection location with flight Mach number, used in combination with a diverging area combustor, provides effective throat area control and greatly simplifies combustor design issues. Because of high combustor pressure and temperature, minimization of combustor length is important to overall engine weight. Combustor cooling is also complicated by continual variation in heating patterns as shock structure changes with flight Mach number and throttle setting. A related scramjet design issue is thermal balance of the cooling and engine thrust requirement on fuel use. Thermally efficient structural design of cooling jackets can significantly reduce cooling requirements, and thus minimum fuel use and engine specific impulse.
The end of the combustion process, where the fuel completes burning, is effectively the beginning of the nozzle. The ideal nozzle arrangement will allow the flow to expand gradually and turn along the base of the vehicle until it reaches the pressure corresponding to the atmosphere and its direction is parallel to the direction of thrust. The typical nozzle will consist of two components, the base of the vehicle itself, and a lower flap which is adjustable to correct for variations in flight conditions.
Hypersonic airbreathing engines can operate on a variety of fuels, including hydrogen and hydrocarbons. Liquid hydrogen is the choice for space-launched vehicles because of its huge heat capacity, which is used regeneratively to cool the engine and vehicle before being burned in the engine. The lower heat capacity of hydrocarbon fuels limits their application to less than Mach 8. At higher Mach numbers using hydrocarbon fuel, more fuel is required to cool the vehicle than is used for propulsion, resulting in wasted fuel.
There are a series of important technical issues facing any scramjet designer. Some of them include achieving rapid and efficient fuel mixing and combustion, effective inlet combustor isolation, limiting peak heating rates (shock interactions with the engine leading edges), balancing total cooling requirements with fuel heat capacity and engine fuel usage rate, and integrating with the vehicle airframe and low-speed systems. Finally, it must be added that the Earth-to-orbit mission is one that is so technically challenging that any advances in high temperature, lightweight materials must be used and in fact will greatly enhance mission success.
There is another class of nontraditional engines that shows great promise. These are the hybrids. These engines are composed of solid fuel grains, usually in annular form, similar to that in solid rocket motors but with no oxidizer added to the slurry before it is poured. Instead, a liquid oxidizer is used and is introduced by having it flow (in gaseous form) through the annular region. Combustion occurs at the interface. This system has many important benefits. Unlike traditional solids, thrust can be throttled or terminated simply by varying the flow of oxidizer. It is extremely easy to cast and handle because the fuel by itself can be made fairly inert. This leads to a relatively safe system. Its performance is between those of solids and liquids. On the other hand, as motor size increases, usually by lengthening the grain, efficient combustion becomes harder to achieve because of the difficulty of providing a sufficient amount of unburned oxidizer at the gas-solid boundary. To date, several small experimental motors have been successful, but no motors in the Earth-to-orbit class have been demonstrated. With improvements in understanding the physics of mixing, it is felt that the problems of hybrids will eventually be overcome.
Summary of Design Process
The overall design requirements for the propulsion system flow from the vehicle system. The vehicle system’s design requirements are derived from the mission or missions that have been specified technically and programmatically. The mission may specify that the vehicle must deliver a payload to an orbital point (i.e., LEO—low Earth orbit or GTO—geosynchronous transfer orbit). The mission could also be required to attain a Mach number at a certain altitude. All of these requirements indicate that the vehicle will have to be propelled by some type of propulsion system that must directly interface with the vehicle and will explicitly affect the propellant fraction. Additionally, the mission architecture will indicate whether the vehicle system is reusable or expendable. The flight rate will greatly affect the way the architecture is employed and the way the overall system life-cycle cost is determined. Therefore, the vehicle and mission architecture can heavily influence the development, production, and operating cost of the propulsion system as well.
Determining the type of propellant or rocket engine cycle that is optimum to maximize the mission and the cost-effectiveness of the overall vehicle system requires what is commonly known as multidisciplinary optimization or hyper-functional integration/optimization. All of the design choices that influence the propulsion system (e.g., cycle, propellants and ratio, chamber pressure, nozzle area ratio) directly affect the design thrust size and specific impulse. They also affect the cost of the propulsion system. Those propulsion design parameters in turn influence the vehicle size and/or payload delivery capability. Thus, many propulsion and vehicle design variables must be simultaneously or interactively examined to arrive at the optimum design. Here, the primary propulsion variables that interact with the vehicle “physics” are the thrust size, engine thrust-to-weight, and the engine specific impulse or Isp (vacuum and sea level, or trajectory average). Systems analysis of the integrated design problem will consider vehicle design parameters, such as vehicle thrust-to-weight, wing loading (if winged), maximum flight path thermal and structural loading (i.e., maximum dynamic pressure), and vehicle volumetric efficiency (i.e., bulk density). These vehicle design parameters will be optimized with variations on the propulsion design parameters such as chamber pressure, nozzle area ratio, and propellant ratio. If the vehicle is designed for horizontal takeoff and ascent, then the design parameters will include inlet design flow capture as well as propulsion system integration with the airframe.
Selection of the “best” vehicle and propulsion system design will be determined by which combination of design variables provides the lowest cost, highest mission performance per unit cost, and lowest complexity and risk.