This chapter covers a general presentation of the process of selecting rocket propulsion systems for a given mission. There are many factors, restraints, and analyses that need to be considered and evaluated before suitable selections can be properly made. The objectives are to operate at very high combustion efficiencies and prevent recurring or destructive combustion instabilities. See Refs. 19–1 to 19–6. Because design problems most often have several possible engineering solutions, this can be a complicated task. Three specific aspects of selection are covered here in some detail:

1. Comparison of merits and disadvantages of liquid propellant rocket engines with solid propellant rocket motors and hybrids.
2. Key factors used in evaluating particular rocket propulsion systems and in selecting from among several competing candidates.
3. Necessary interfaces between the propulsion system and the flight vehicle and/or the overall system.

A propulsion system is really one of several subsystems of a flight vehicle. The vehicle, in turn, can be part of a larger overall system. An example of an overall system would be a communications network with ground stations, transmitters, and several satellites; each satellite may require a multi‐stage flight vehicle during launch with all associated propulsion systems as well as on‐orbit attitude‐control propulsion systems (each with specific propulsion requirements). Additionally, the length of active time in orbit is a system parameter that affects both satellite size and the total impulse propulsion requirements.

Vehicle subsystems (such as its structure, power supplies, propulsion, guidance, control, communications, ground support, and/or thermal control) often pose conflicting requirements. Only through careful analyses and proper system‐engineering studies is it possible to find compromises that allow all subsystems to operate satisfactorily and in harmony with each other. The “systems engineering” approach is now utilized routinely, see Ref. 19–1, and engineering design has advanced considerably in recent times with computer‐aided design (CAD) now commonly used. Certain publications address specifically the design of space systems (e.g., Refs. 19–1 to 19–3) and the design of liquid propellant engines (e.g., Ref. 19–4). In our book, Sections 11.6, “System Integration and Engine Optimization,” and 15–4, “Rocket Motor Design Approach,” are preludes to this chapter and may contain some duplicate content.

Systems engineering is a useful discipline for rocket propulsion systems selection. This topic can be defined in several ways; one being (adapted from Ref. 19–5) “a logical process of activities, analyses and engineering designs, that transforms a set of requirements arising from specific mission objectives in an optimum way. It ensures that all likely aspects of a project or engineering system have been considered and integrated into a consistent whole.” Such studies comprise all elements of a propulsion system and its ground support, all interfaces with other vehicle subsystems or ground‐based equipment, and consider safety aspects, risks, and costs, together with the orderly selection of the most suitable propulsion system.

There are commonly three levels of requirements from which propulsion system requirements can be derived, (see Ref. 19–5). At each level, certain activities, studies, and/or trade‐off evaluations are involved.

1. At the top level are mission defining requirements, such as space communications or missile defense. Here, analyses and optimizations are usually carried out by the mission responsible organizations. They define rocket propulsion parameters like trajectories, orbits, payloads, number of vehicles, life, and the like, and document them as mission requirements specifications.
2. From the mission requirements document, definitions and specifications for the flight vehicle are derived. These may include vehicle size, vehicle mass, number and size of vehicle stages, types of propulsion, required mass fraction, minimum specific impulse, thrust vector and/or attitude control needs, allowable accelerations, acceptable propellants, restraints on engine size and/or inert mass, cost restraints, allowable start delays, desired thrust–time profiles, preferred propulsion system types (liquid, solid, hybrid, electrical), ambient temperature limits, and so forth. Engineering analyses together with preliminary designs and additional specifications are usually done by the organization responsible for the flight vehicle.
3. Propulsion requirements form the basis for the final design and development of the propulsion system, and they are derived from the above two levels. These may include, for example, an optimum thrust–time profile, restart requirements, liquid propellant engine cycle, pulsing schedules, solid rocket motor case L/D (length/diameter), grain configuration, restraints on travel of the center of mass, details of thrust vector and/or attitude control mechanisms, reliability, cost and/or schedule restraints, and expected number of propulsion systems and spare parts to be delivered. These requirements are often done by the vehicle development organization with inputs from one or more selected propulsion organizations.

Much of the design (i.e., three‐dimensional modeling of the propulsion system), most of the analyses (i.e., stress analysis), testing, keeping track of all hardware assemblies and part inventories, and inspection records or NC programming are done today using computers. Interfacing these computer programs with each other and with organizational sectors that deal with some or all aspects of the propulsion system becomes an important objective.

All mission, vehicle, and propulsion system requirements can be related to performance, cost, and/or reliability. In general, if cost criteria have the highest priority then the propulsion system will be different from one that has performance as its top priority. For any given mission, one criterion is usually more important than the others. There is a strong interdependence between the three levels of requirements and the criteria categories mentioned. Some propulsion system requirements (which is usually a second‐tier subsystem) strongly influence the vehicle and vice versa. An improvement in propulsion performance, for example, can affect vehicle size, overall system cost, and/or life (which normally translate into reliability and cost).

19.1 SELECTION PROCESS

The selection process for the vehicle and its associated rocket propulsion systems is an integral part of any overall vehicle design effort. Selection is based on a series of criteria largely derived from the requirements, which serve to evaluate and compare alternate propulsion system proposals. In addition to the chosen application, determining the most suitable rocket propulsion system depends on the experience of those making the selection and their ability to express many propulsion system characteristics quantitatively, the quality and amount of applicable data available, and on available time and resources for examining alternate propulsion systems. We describe here a somewhat idealized selection process, the one depicted in Fig. 19–1, but there are other sequences to do the job of selection.

Because total vehicle's performance, flight control, operation, and/or maintenance critically depend on the performance, control, and operation of the rocket propulsion system (and vice versa), the selection process will normally consist of several iterations in defining both the vehicle and propulsion requirements that satisfy the given mission. This iterative process involves both the system's organization (and vehicle/system contractor) and one or more propulsion organizations (or rocket propulsion contractors). Documentation and reporting can take many forms; electronic storage has greatly expanded capabilities to network, record, and retrieve documents.

Several competing candidate systems are usually evaluated. They may originate from different rocket propulsion organizations, perhaps on the basis of modifications of some existing rocket propulsion system, or may include some novel technology, or may be new types of systems specifically configured to fit the vehicle or mission needs. In making evaluations it will be necessary to compare the candidate propulsion systems with each other and to rank‐order them (in accordance with some selection criteria on how well they meet each requirement). This necessitates studies for each candidate system and also, sometimes, additional testing. For example, statistical analyses of the functions, failure modes, and safety factors of all key components may lead to quantifiable reliability estimates. For some criteria (such as safety or prior related experience) it may not be possible to compare candidate systems quantitatively but only subjectively.

For any given mission, several rocket parameters often need to be optimized. Trade‐off studies can be used to optimize these parameters in accordance with overall requirements. Such trade‐off studies have often been performed on the following items: number of thrust chambers, optimum chamber pressure (considering performance and inert propulsion system mass), selected mixture ratio (performance, tank volume), solid propellant case configuration (L/D), cost versus propulsion system performance, chamber pressure versus size and/or mass of propulsion system, chamber pressure versus heat transfer or cooling method, alternate grain configurations, minimum center of mass travel during operation, thrust magnitude/profile versus powered flight duration, nozzle exit area ratio versus performance, nozzle mass, and/or nozzle dimensions (length, diameter), alternate TVC (thrust vector control) concepts versus mass and power, chamber pressure versus heat transfer (cooling method or insulation) and inert mass. See Ref. 19–5.

Early in the selection process a tentative recommendation is made as to whether the propulsion system should be a solid propellant motor, a liquid propellant engine, an electrical thruster, a solar‐thermal system, or some combination of these. Each type has its own thrust regime, specific impulse, thrust‐to‐weight ratio (acceleration), and/or likely duration, as indicated in Table 2–1 and Fig. 2–4, where these factors have been listed for several chemical rocket propulsion devices and several types of nonchemical engines. Liquid engines and solid motors are covered in Chapters 6 to 15, hybrids in Chapter 16, and electric propulsion in Chapter 17.

If an existing vehicle is to be upgraded or modified, the question arises if its propulsion system should also be changed (e.g., higher thrust, more total impulse, or faster thrust vector control). While there might be some trade‐off studies leaning toward modifying propulsion parameters, in upgrades one normally does not consider an entirely different propulsion system, as is often done in entirely new vehicles or missions. Also, it is rare that an identical rocket propulsion system will satisfy two different applications; frequently, design changes, interface modifications, and requalification will be necessary to adapt an existing rocket propulsion system to another application. Most often, proven qualified propulsion systems that fit the desired requirements have an advantage in cost, scheduling, and reliability.

Electric propulsion systems address a set of unique space applications with low thrusts, low acceleration trajectories, high specific impulses, long operating times, and relatively massive power supplies. They perform competitively for certain space transfer and orbit maintenance missions. With more electric propulsion systems flying than ever before, the choice of proven electric propulsion thruster types has widened. At the time of this writing, there are plans for all‐electric‐propulsion upper launch stages. These systems, together with their design approaches, are described in Chapter 17.

When chemical propulsion systems are deemed most suitable for a particular application, selection has to be made between liquid propellant engines, solid propellant motors, or hybrid rocket propulsion systems. Some of the major advantages and disadvantages of liquid propellant engines and solid propellant motors are presented in Tables 19–1 through 19–4. These lists may not be complete, some items need further information, some items may be controversial, and a number are restricted to particular applications. Most of the entries from these lists may be converted to evaluation criteria; for any specific mission, relevant entries would be rank‐ordered in accordance with their relative importance. A quantification of many of these items would be needed.

The question is often asked: which are better, solid or liquid propellant rocket propulsion systems? A clear statement of strongly favoring one or the other can only be made when referring to a specific set of flight vehicle missions. Today, solid propellant motors (SPRMs) seem to be preferred for tactical missiles (air‐to‐air, air‐to‐surface, surface‐to‐air, or short‐range surface‐to‐surface) and ballistic missiles (long‐ and short‐range surface‐to‐surface) because of their instant readiness and compactness; the absence of spills or leaks of hazardous liquids are also important criteria in these applications. Liquid propellant rocket engines (LPREs) seem to be preferred for space‐launched main propulsion units and upper stages, because of their higher specific impulse, relatively clean exhaust gases, and random throttling capability. They are also favored for postboost control systems and attitude control systems, because of their random multiple pulsing capabilities with precise impulse cutoff, and for pulsed axial and lateral thrust propulsion on hit‐to‐kill interceptors. There are, however, always exceptions to these preferences.

When selecting rocket propulsion systems for a major new multiyear high‐cost project, considerable time and effort are spent in evaluation and in developing rationales for quantitative comparisons. In part, this is in response to government policies as well as to international competition. Multiple studies are often done by competing systems and rocket propulsion organizations; formal reviews are then used to assist in arriving at a proper selection.

When submitting a proposal for a new or modified unit, it is necessary to have a well‐defined project plan, a proper cost estimate, and a realistic schedule for the proposed work. These can only be prepared after the following become available: (1) design layouts of the geometry with functions and features of the proposed system, (2) descriptions of planned tests, data handling, special test equipment, and safe test facilities, (3) plans for manufacturing that include fabrication steps, tooling fixtures, available suitable factory space, and special manufacturing equipment, (4) lists of qualified parts suppliers and other vendors, (5) descriptions and availability of all needed materials, and (6) a list of experienced personnel.

19.2 CRITERIA FOR SELECTION

Criteria used in selecting a particular rocket propulsion system are unique to the particular mission or vehicle application. However, some selection factors may apply to several applications. Typical performance criteria include propellant composition, thrust profile, minimum specific impulse, operating duration, thrust variation (% throttling), life, system mass or weight, nozzle exit area ratio, solid propellant parameters (strength, elongation, or storage temperature limits), maximum allowed residual liquid propellant or solid propellant grain slivers, and ignition parameters; for reusable vehicles these criteria should include number of restarts, provisions for draining residual liquid propellants and for cleaning and drying out tank feed systems. Most tactical missiles with SPRMs include criteria to account for tolerating external fires, bullet impacts, or nearby explosions that do not cause propellant detonation. For storage or transport over public conveyances, rules and safety standards regarding rocket propulsion systems loaded with solid propellants must be well understood. Allowable tolerances/variations for each of the items above need to be well defined. For a more complete list of these criteria see Ref. 19–6.

Any evaluation/selection groups will include experienced technical personnel from the vehicle development/fabrication organization, and such groups might be augmented by personnel from Government Laboratories, consultants, and/or other outside experts. Proposal selection is done in accordance with well‐established and prioritized criteria by personnel from the organization that will do the contracting of the propulsion system. Actual selection will depend on the balancing of various relevant factors in accordance with their importance, benefits, and/or potential impact on the system, and on quantifying as many of these selection factors as possible through analysis, extrapolation of prior experience/data, cost estimates, weights, and/or separate tests. See Ref. 19–5. Layouts, inert mass estimates, center‐of‐gravity analyses, vendor cost estimates, preliminary stress or thermal analyses, and other preliminary design efforts are necessary to quantify the selection parameters. A comparative examination of the interfaces of alternate propulsion systems can also be a part of the process.

For a spacecraft that contains optical instruments (e.g., telescopes, horizon seekers, star trackers, or infrared radiation seekers) the exhaust plume must be free of possible contaminants that may deposit or condense on optical windows, mirrors, lenses, photovoltaic cells, or radiators, and degrade their performance, and also be free of particulates that scatter sunlight into the aperture of optical instruments, thus causing erroneous signals. Here, conventional composite solid propellants and pulsing storable bipropellants are not satisfactory, but cold or heated clean gas jets (H₂, Ar, N₂, etc.) and monopropellant hydrazine reaction gases are usually acceptable. This topic is discussed in Chapter 20. Another example is the need for smokeless propellant‐exhaust plumes for avoiding any visual detection of smoke or vapor trails. This applies particularly to tactical missiles. Only a few solid propellants and several liquid propellants can be nearly smokeless, essentially free of vapor trails under most weather conditions.

Sometimes selection criteria conflict with each other. For example, some propellants with a very high specific impulse are more likely to experience combustion instabilities; higher chamber pressures increase the pressure ratio across the nozzle and will give an increase in specific impulse, improving vehicle performance, but the combustion chamber (and in LPREs also the propellant feed system) will require thicker, heavier walls and this inert mass increase will reduce the vehicle's mass ratio, which will in turn reduce the vehicle velocity (see Sec. 4–6); in electric propulsion, their high specific impulse is usually accompanied by low thrust and massive power‐generating and conditioning equipment. Compromises must to be made when propulsion requirements are incompatible with each other. For example, the monitoring by extra sensors can prevent occurrence of certain types of failures by promoting remedial action and thus enhancing propulsion system reliability, yet these extra sensors and control components contribute to system mass and complexity, and their possible failure can reduce overall reliability. The selection process may also include feedbacks for when the stated propulsion requirements cannot be met or do not make sense, that leads to revisions or redefinitions of initial mission or vehicle requirements.

Once cost, performance, and reliability drivers have been identified and quantified, priorities for the system's criteria can be identified and the selection of the best propulsion system for a specified mission can proceed. The final propulsion requirements may result after several iterations take place, and will usually be documented, for example, in a propulsion requirement specification. A substantial number of records will be required here (such as rocket engine or rocket motor acceptance documents, CAD (computer‐aided design) images, parts lists, inspection records, laboratory test data, etc.). Usually, there are many specifications associated with design and manufacturing as well as with vendors, materials, and so on. There must also be disciplined procedures for making and approving design and manufacturing changes. This stage now becomes the starting point for the design and development of the selected propulsion system.

19.3 INTERFACES

An interface can be considered to be a surface (usually irregular) forming the boundary between two adjacent bodies or assemblies. Three interfaces types are of concern to rocket propulsion systems (RPSs). The first interface is between the system and its vehicle—these two are usually designed and fabricated separately but must form when assembled a single viable structure which transmits (at the interface) electrical power and/or electrical signals (e.g., hydrazine heaters, command and control/start signals, measuring instruments electric outputs, TVC power/signals, etc.), and in some cases transfer fluids (pressurized gas and loading liquid propellants). For large vehicles, a second interface exists between the RPS and the launch stand on the ground; here the same interface connections apply as above, but there are additional provisions to allow separation of the vehicle from the launch stand at liftoff. This includes explosive bolts to release the vehicle structurally from the launch stand, electrical disconnect devices (often wire cutters), and fluid disconnect fittings (valves) in the fill pipes designed for minimum spill losses of gas or liquid propellant. The third interface is between any two stages in a multistage flight vehicle; for stage separation all structural interconnect links and all electric connections between the two stages need to be disconnected or severed just prior to separation. This may include using any remaining high‐pressure gas (or small separation rocket motors) for reverse thrust on the lower stage and these actions need to be controlled during the separation maneuvers.

In Section 19.2 interfaces between the propulsion system and the vehicle and/or overall system were identified as criteria to be considered for study in the selection of propulsion systems. See also Ref. 19–5. Only a few rocket propulsion systems are straightforward to integrate and interface with their vehicles. Furthermore, interfaces are an important aspect of any design and development discipline. Interfaces assure system functionality and compatibility between the propulsion system and the vehicle's other subsystems under all likely operating conditions and mission options. Usually, an interface document or specification is prepared for use by designers and operating and/or maintenance personnel.

A list of systems in increasing order of complexity follows. Along with cold gas systems, simple solid propellant rocket motors have the fewest and least complex set of interfaces. Monopropellant liquid rocket engines also have relatively few and simple interfaces. Solid propellant motors with TVC and thrust termination capabilities have additional interfaces compared to simple motors. Bipropellant rocket engines are more complex than monopropellant LPREs. The number and difficulty of interfaces increase when they have turbopump feed systems, throttling features, TVC, and/or pulsing capabilities. In electric propulsion systems the number and complexity of interfaces is highest for electrostatic thrusters with pulsing capabilities, when compared to steady‐state electrothermal systems; in general, the more complex electric propulsion systems produce higher values of specific impulse. When the mission includes recovery and reuse of the propulsion system or in manned vehicles (where the crew can monitor and/or override propulsion system commands), several more additional interfaces, safety features, and requirements need to be introduced.

19.4 COST REDUCTION

For the past 15 years there has been increased emphasis on cost reduction in rocket propulsion systems for space flight. Table 19–5 lists some cost reduction examples that have been achieved to date. This list is neither complete nor comprehensive and only some items may apply to any one particular system, but it does represent many relevant issues. A reduction in cost is acceptable only when it does not diminish or compromise the intended function or flight performance or environmental compatibility of the affected components, subassemblies or systems; it must always allow the vehicle to fly and complete the intended mission within all its overall constraints.

REFERENCES

1. 19–1. W. J. Larson and J. R. Wertz (Eds.), Space Mission Analysis and Design, 2nd ed., Microcosm, Inc., and Kluver Academic Publishers, Boston, 1992.
2. 19–2. J. C. Blair and R. S. Ryan, “Role of Criteria in Design and Management of Space Systems”, Journal of Spacecraft and Rockets, Vol. 31, No. 2, Mar.–Apr. 1994, pp. 323–329.
3. 19–3. R. W. Humble, G. N. Henry, and W. J. Larson, Space Propulsion Analysis and Design, McGraw‐Hill, New York, 1995.
4. 19–4. D. K. Huzel and D. H. Huang, Modern Engineering for Design of Liquid Propellant Rocket Engines, rev. ed. Progress in Astronautics and Aeronautics, Vol. 147, AIAA, Washington, DC, 1992.
5. 19–5. P. Fortescue, J. Stark, and G. Swinerd, Spacecraft Systems Engineering, 3rd ed., John Wiley & Sons, Hoboken, NJ, 2003.
6. 19–6. Table 19–6 (pgs. 703 to 705) and Table 9–7 (pgs. 706 to 708) of G. P. Sutton and O. Biblarz, Rocket Propulsion Elements, 8th Ed., John Wiley & Sons, Hoboken, NJ, 2010.
7. 19–7. S. Schmidt et al., “Advanced ceramic matrix composite materials for current and future propulsion technology applications,” Acta Astronautica, Vol. 55, Nos. 3–9, Aug.–Nov. 2004, pp. 409–420.