Economic Factors

Availability in quantity and a low cost are very important considerations in propellant selection. In military applications, consideration has to be given to the logistics of production, supply, storage, along with other factors. The production process should require only ordinarily available chemical equipment and available raw materials. It is more expensive to use toxic or cryogenic propellants than storable, nontoxic ones, because the former require additional steps in their operation, more safety provisions, additional design features, longer check‐out procedures prior to launch, and often better trained personnel.

Performance of Propellants

The performance rocket engines may be compared on the basis of their specific impulse, exhaust velocity, characteristic velocity, and/or other engine parameters. These were introduced in Chapters 3, 5, and 6. The specific impulse and exhaust velocity are functions of pressure ratio, specific heat ratio, combustion temperature, mixture ratio, and molecular mass. Equilibrium values of performance parameters for various propellant combinations can be calculated with a high degree of accuracy and several are listed in Table 5–5. Very often, performance is also expressed in terms of flight performance parameters for specific rocket applications, as explained in Chapter 4. Here, average propellant density, total impulse, and engine mass ratio usually enter into the various flight relation descriptions.

For high performance, high chemical energy content per unit of propellant mixture is desirable because this yields high chamber temperatures. A low molecular mass for the combustion products gases is also desirable. This can be accomplished by using fuels rich in hydrogen obtained when a significant portion of the hydrogen gas injected or produced remains uncombined. In general, therefore, the best mixture ratio for many bipropellants is not stoichiometric (which results in complete oxidation and yields the highest flame temperature) but fuel‐rich, containing a large amounts of low‐molecular‐mass reaction products as shown in Chapter 5.

When relatively small metallic fuel particles (such as beryllium or aluminum) are suspended in the liquid fuel, it is theoretically possible to increase the specific impulse by between 9 and 18%. A particular chemical propellant combination having the highest ideal specific impulse known (approximately 480 sec at 1000 psia chamber pressure and expansion to sea‐level atmosphere, and 565 sec in a vacuum with a nozzle area ratio of 50) uses a toxic liquid fluorine oxidizer with hydrogen fuel plus suspended toxic solid particles of beryllium; as yet, acceptably safe and practical means for storing such seeded propellants and their practical rocket engine use remain undeveloped.

Gelled propellants are materials that have additives that make them thixotropic. They have the consistency of thick paint or jelly when at rest, but do liquefy and flow through pipes, valves, pumps and/or injectors when an adequate pressure or shear stress is applied. In spite of extensive research and development work and demonstrations of better safety and certain “green” qualities, to date they have not been adopted for any production rocket engine. Gelled propellants have been described in the Sixth, Seventh, and Eighth editions of this book.

Common Physical Hazards

Although several hazard categories are described below, they do not all apply to each propellant or to every bipropellant combination. Hazards can be different for each specific propellant and must be carefully understood before working with it. The consequences of unsafe operation or unsafe design are usually also unique to each propellant.

Desirable Physical Properties

High Specific Gravity

Denser propellants provide a larger propellant mass for a given vehicle tank volume. Alternatively, for a given mass, they permit smaller tank volumes and, consequently, lower structural vehicle mass and lower aerodynamic drag. The specific gravity of a propellant, therefore, has an important effect on the maximum flight velocity and range of any rocket‐powered vehicle or missile flying within the Earth's atmosphere as explained in Chapter 4. Specific gravities for various propellants are plotted in Fig. 7–1. Variations of ambient temperature in stored propellants cause changes of the liquid level in their storage tanks.

For a given bipropellant mixture ratio r, the average specific gravity of any propellant combination δ_av can be determined from the specific gravities of the fuel δ_f and of the oxidizer δ_o. This average specific gravity δ_av is defined below. Here r is the mixture ratio; it represents the oxidizer mass flow rate divided by the fuel mass flow rate (see Eq. 6–1):

7–1

Values of δ_av for various propellant combinations are listed in Table 5–5. The value of δ_av can be increased by adding high density materials to the propellants, either by solution or colloidal suspension. An identical equation can be written for the average density ρ_av in terms of the fuel and oxidizer densities:

7–2

Though the specific gravity is unitless, in the SI system it has the same numerical value as the density expressed in units of grams per cubic centimeter or kg/liter. In some performance comparisons the parameter density specific impulse I_d is used. It is defined as the product of the average specific gravity δ_av and the specific impulse I_s:

7–3

A propellant density increase will allow increases in mass flow and total propellant mass when all other system parameters remain the same. For example, Fig. 7–1 shows that lowering the temperature of liquid oxygen from –250 to –280 °F raises its specific gravity by approximately 8%. Therefore, the system's mass flow rate and the total mass will increase by approximately the same amount. These changes will enable increases in chamber pressure, total impulse, and thrust; changes of less than 1% in specific impulse may also be noticed. So there is some benefit to vehicle performance in operating with the liquid propellants at their lowest practical temperature. In ground‐based systems, cooling of liquid oxygen may be achieved with a (colder) liquid nitrogen heat exchanger just before launch. The percent increase in performance with propellants other than oxygen is usually much smaller.

Ignition, Combustion, and Flame Properties

When a propellant combination is spontaneously ignitable burning is initiated as soon as the oxidizer and the fuel come in contact with each other. Spontaneously or self‐ignitable propellant combinations are called hypergolic propellants. Although ignition systems are not necessarily objectionable, their elimination simplifies the propulsion system. In order to reduce potential explosion hazards during starting, all rocket propellants should be readily ignitable and exhibit acceptably short ignition time delays in order to eliminate potential explosion hazards. Starting and ignition problems are discussed further in Section 8.6.

Nonspontaneously ignitable propellants must be energized by external means for ignition to begin. Igniters are devices that accomplish a localized initial chamber pressurization and initial heating of the propellant mixture to a state where steady flow combustion can be self‐sustained. The amount of energy needed from the igniter to activate the propellants should be small so that low‐power, light‐weight ignition systems may be used. The energy required for satisfactory ignition usually diminishes with increasing propellant storage temperature. At low ambient temperatures ignition can become relatively slow (0.05 to 0.02 sec).

Certain propellant combinations burn very smoothly, without vibration (i.e., gas pressure oscillations). Other propellant combinations do not exhibit such combustion stability and, therefore, become less desirable. Combustion stability is treated in Chapter 9.

Smoke formation is objectionable in many applications because it may deposit on the surrounding equipment and parts. Smoke and brilliantly luminous exhaust flames are objectionable for certain military applications since they can be easily detected. In some applications, the condensed species from gaseous exhausts can cause surface contamination on spacecraft windows or optical lenses, and the presence of free electrons in a flame may cause undesirable interference or attenuation of communications radio signals. See Chapter 20 for information on exhaust plumes.

Property Variations and Specifications

Propellant properties and quality must not vary from delivered batch to batch because this can affect engine performance and combustion from changes in physical and/or chemical properties. The same propellant must have consistent composition and storage properties; rocket operating characteristics should not change even when propellants are manufactured at different times or made by different manufacturers. For these reasons propellants are purchased to conform to strict specifications which define ingredients, maximum allowable impurities, packaging methods and/or compatible materials, allowable tolerances on physical properties (such as density, boiling point, freezing point, viscosity, or vapor pressure), quality control requirements, container cleaning procedures, documents of inspections, laboratory analyses, and/or test results. A careful chemical analysis of composition and impurities is always necessary. Reference 7–7 describes some of these methods of analysis.

Additives

Altering and tailoring propellant properties may be achieved with additives. For example, a reactive ingredient is added to make a nonhypergolic fuel become hypergolic (readily ignitable). To desensitize concentrated hydrogen peroxide and reduce self‐decomposition, it is diluted with 3 to 15% of pure water. To increase the density or to alleviate certain combustion instabilities, a fine powder of a heavy solid material (such as aluminum) is suspended in the fuel. The use of additives to lower the freezing‐point temperature of nitrogen tetroxide is treated later.

Liquid Oxygen (O₂) (LOX)

Liquid oxygen is listed in Table 7–1. It is widely used as an oxidizer and burns with a bright white‐yellow flame with most hydrocarbon fuels. It has been used in combination with alcohols, jet fuels (kerosene‐type), gasoline, and hydrogen. As shown in Table 5–5, attainable performances are relatively high, and LOX is therefore a desirable and commonly used propellant in large rocket engines. The following missiles and space launch vehicles burn oxygen: (1) with jet fuel or RP‐1 (kerosene)—Atlas V and Soyuz (Russia); (2) with hydrogen—Ariane‐V (France), Delta IV, and Centaur upper stage; (3) with ethanol—the historic V‐2 (German) and the Redstone missile. Figs. 1–12, 1–13, and 6–1 show units that use LOX; Figs. 5–1 to 5–6 present theoretical performance information for LOX with one kerosene‐type fuel.

Although it usually does not burn spontaneously with organic materials at ambient pressures, combustion and/or explosions have occurred when confined mixtures of oxygen and organic matter suddenly come under pressure. Impact tests show that mixtures of LOX with many commercial oils or organic materials will detonate. Liquid oxygen also supports and accelerates the combustion of other materials. Handling and storage can be safe only when all contact materials are clean. Liquid oxygen is a noncorrosive and nontoxic liquid that will not cause the deterioration of clean container walls. Because of its low temperature, this propellant causes severe frostbite during any prolonged contact with human skin. Because LOX evaporates rapidly at ambient conditions, it cannot be stored readily for any extended length of time. When LOX is used in large quantities, it is often produced very close to its geographical point of application. Liquid oxygen can be obtained in several ways, but most commonly from the fractionated distillation of liquid nitrogen out of liquid air.

It is necessary to insulate all lines, tanks, valves, and parts that contain liquid oxygen in order to reduce evaporation losses. Rocket propulsion systems which have to remain filled with liquid oxygen for several hours and liquid oxygen storage systems need to be well insulated against heat absorption from the surroundings. External drainage provisions have to be made on all LOX tanks and lines to allow the water that condenses on the cold outside walls to drain from the rocket during launch preparations. Provisions for minimizing and removing ice formations are also needed.

Hydrogen Peroxide (H₂O₂)

This propellant is not only a powerful liquid oxidizer when burning with an organic fuel but also clean burning. It produces a nontoxic exhaust when used as a monopropellant. In rocket applications, hydrogen peroxide needs to be in a highly concentrated form of 70 to 98% (known as high‐test peroxide or HTP), the remainder being mostly water. Hydrogen peroxide was used in gas generators and rocket applications between 1938 and 1965 (X‐1 and X‐15 research aircraft). Since that time, 90% hydrogen peroxide has been the most common concentration for rocket engine use and gas generator applications.

As a monopropellant, it decomposes according to the following chemical reaction, forming superheated steam and hot gaseous oxygen:

This decomposition is brought about by the action of catalysts such as silver screens, various liquid permanganates, solid manganese dioxide, platinum, or iron oxide but, in fact, many materials can catalyze hydrogen peroxide. Theoretical specific impulses for 90% hydrogen peroxide can be around 154 sec, when used as a monopropellant with a solid catalyst bed. As bipropellant, H₂O₂ is hypergolic with hydrazine and will burn well with kerosene.

Concentrated peroxide causes severe burns when in contact with human skin and may ignite when in contact with wood, oils, and many other organic materials. In the past, rocket engines with hydrogen peroxide as oxidizer have been used for aircraft boosting (German Me 163, and U.S. F 104) and for satellite launching (Britain: Black Arrow). As a monopropellant it has been used for attitude control thrusters in the Soyuz space capsule and for steam generators driving the main propellant pumps in the Soyuz launch vehicle. It was not used in the United States for some time, partly because it was hard to predict its long‐term storage stability. In the past, concentrated hydrogen peroxide would self‐decompose during storage at about 1% per year. This progressively diluted this propellant until no longer useful. As oxygen gas would bubble out of the stored liquid, the decomposition rate would accelerate. Small amounts of impurities in the liquid or in wall materials would also accelerate decomposition and cause a rise in liquid temperature. Before reaching 448 K (175 °C), warm stored H₂O₂ or contaminated H₂O₂ must be diluted and discarded because an explosion may occur.

Recent progress in the manufacture of storage tanks and piping materials (and stricter cleaning methods) has much reduced the amount of impurities and lengthened storage life of HTP from about 3 to 4 years to 12 to 16 years, and there may further improvements. This has much renewed interest in this dense oxidizer and several different investigations and development programs have been underway to date (Ref. 7–8). To date, the authors have not become aware of any new production applications for 90% H₂O₂.

Nitric Acid (HNO₃)

Several types of nitric acid mixtures were used as oxidizers between 1940 and 1965; these are not used extensively today in the U.S. The most common type, red fuming nitric acid (RFNA), consists of concentrated nitric acid (HNO₃) that contains between 5 and 27% dissolved nitrogen dioxide (NO₂). Compared to concentrated nitric acid (also called white fuming nitric acid), RFNA is more energetic, more stable in storage, and slightly less corrosive to many tank materials.

Only certain types of stainless steel, gold, and a few other materials are satisfactory as storage containers and/or tubing materials for concentrated nitric acid. Small additions of fluoride ion (less than 1% of hydrofluoric acid, or HF) inhibit nitric acid corrosion, forming a protective fluoride layer on walls and reducing the corrosion with many metals. This combination is called inhibited red fuming nitric acid (IRFNA). Even with an inhibitor, nitric acid reacts with many wall materials and forms dissolved nitrates and sometime insoluble nitrates. This changes the properties of this oxidizer and may cause blocking of valve and injector orifices. Nitric acid reacts with gasoline, various amines, hydrazine, dimethylhydrazine, and alcohols. It ignites spontaneously with hydrazine, furfuryl alcohol, aniline, and other amines. The specific gravity of nitric acid varies from 1.5 to 1.6, depending on the percentages of nitrogen dioxide, water, and impurities. This high density permits compact vehicle construction.

When RFNA is exposed to air, the evaporating red‐brown fumes are exceedingly annoying and poisonous. In case of accidental spilling, the acid must be quickly diluted with water or chemically deactivated. Lime and alkali metal hydroxides and carbonates are common neutralizing agents. However, any nitrates formed by the neutralization also act as oxidizing agents and must be handled accordingly. Vapors from nitric acid or red fuming nitric acid have an OSHA eight‐hour personnel exposure limit or a threshold work allowance of 2 ppm (parts per million, or about 5 mg/m³) and a short‐term exposure limit of 4 ppm. Droplets on the skin cause burns producing sores that do not heal readily.

Nitrogen Tetroxide (N₂O₄) (NTO)

The proper chemical name for N₂O₄ is ‘dinitrogen tetroxide’ as widely used by chemists and the chemical industry, but the rocket industry has always called it nitrogen tetroxide or NTO and this is the designation used in this book. It is a high‐density yellow‐brown liquid (specific gravity of 1.44). NTO is hypergolic with hydrazine, monomethylhydrazine (MMH), and unsymmetrical dimethylhydrazine (UDMH); it is the most common storable oxidizer employed today. Its liquid temperature range is narrow and thus easily accidentally frozen or vaporized. It is only mildly corrosive when pure but forms strong acids when moist or allowed to mix with water, readily absorbing moisture from the air. It can be stored indefinitely in sealed containers made of compatible materials. NTO is hypergolic with many fuels but can also cause spontaneous ignition with many common materials, such as paper, grease, leather, or wood. When it decomposes, the resulting NO₂ fumes are reddish brown and extremely toxic. Because of its high vapor pressure, it must be kept sealed in relatively heavy tanks. The freezing point of N₂O₄ may be lowered (for example, by adding a small amount of nitric oxide or NO) but at the expense of a higher vapor pressure and slightly reduced performance. Such mixtures of NO and N₂O₄ are called mixed oxides of nitrogen (MON) and different grades are available between 2 and 30% by weight NO content. Several MON oxidizers have flown.

Nitrogen tetroxide is used as the oxidizer with UDMH in many Russian engines and in almost all their small thrusters. It was also used with MMH fuel in the Space Shuttle orbital maneuver system and multithruster reaction control system, and is still used in many U.S. spacecraft propulsion systems. For many applications, care has to be taken to avoid freezing the nitrogen tetroxide. The OSHA eight‐hour personnel exposure limit is 5 ppm NO₂ or 9 mg/m³.

Nitrous Oxide (N₂O)

This oxidizer is also known as dinitrogen monoxide and as “laughing gas” (a medical anesthetic), see Refs. 7–9 and 7–10. It is a less potent oxidizer than the other four listed above. At ambient temperature it is not flammable but supports combustion at elevated temperatures. Its handling and safety are discussed in Refs. 7–10 and 7–11. Compared to nitrogen tetroxide (N₂O₄) it is much less toxic, its eight‐hour exposure limit being about 20 times less severe. Various catalysts will cause the decomposition of N₂O into O₂ and N₂, and many materials and some impurities are incompatible with N₂O. Its cryogenic liquid temperature range is very narrow, see Table 7–1 for boiling and freezing points; liquid N₂O is not as cold as LOX. Although somewhat difficult to control, this oxidizer can self‐pressurize itself, avoiding the need for separate helium pressurizing systems.

Nitrous oxide has flown in space for the last 14 years. One principal recent flight application has been that of liquid nitrous oxide with HTPB (hydroxyl‐terminated polybutadiene) as a solid fuel in a hybrid rocket propulsion system (Ref. 7–12) for a suborbital manned space plane, which is discussed in Chapter 16. In a bipropellant combination, it has also been tested as the oxidizer with organic liquid fuels and gaseous monopropellants using a catalyst.

Oxidizer Cleaning Process

The internal surfaces of all newly manufactured components for rocket engine liquid oxidizer systems (pipes, tanks, valves, pumps, seals, injectors, etc.) usually undergo a special cleaning process for removing all traces of organic materials or other impurities that can react with the oxidized and cause bubbles or overpressure. This cleaning includes removal of even minute amounts of grease, oils, machine cuttings, fluids, paints, and small deposits of carbon, plastics, and common dust. The cleaning process may differ for each particular oxidizer and consists of scraping‐off visible deposits, successive multiple flushings and rinsing (with appropriate liquids), followed by drying with clean hot air. After the components are “oxygen clean” they are usually sealed to avoid any contamination prior to assembly or operation.

Hydrocarbon Fuels

Petroleum derivatives encompass a large variety of different hydrocarbon chemicals that can be used as rocket fuels. The most common are types already in use with other engines applications, such as gasoline, kerosene, diesel oil, and turbojet fuel. Their physical properties and chemical composition vary widely with the type of crude oil from which they were refined, with the chemical process used in their production, and with the accuracy of control exercised in their manufacture. Typical values are listed in Table 7–2.

In general, these petroleum fuels form yellow‐white, brilliantly radiating flames and give good performance. They are relatively easy to handle and abundant at relatively low cost. Specifically refined petroleum products particularly suitable as rocket propellants carry the designation of RP‐1 or ‐2 (rocket propellant number 1 or 2). These are basically a kerosene‐type fuel mixture of hydrocarbons with a somewhat narrow range of densities and vapor pressures. Under some conditions, hydrocarbon fuels may form carbon deposits on the inside of cooling passages, impeding heat transfer and raising wall temperatures. Ref. 7–13 indicates that such carbon formations depend on fuel temperature in the cooling jacket, the particular fuel, the rate of heat transfer, and the chamber wall material. Fortunately, RP‐1 is low in olefins and aromatics (the ones likely to cause such solid carbonaceous deposits inside fuel cooling passages). RP‐1 has been used with liquid oxygen in many early rocket engines (see Figs. 5–1 to 5–6 and Ref. 7–14). A very similar kerosene‐type fuel is used in Russia today. In about 2003, RP‐2 was substituted for RP‐1 in U.S. rocket applications; the principal difference is a reduced sulphur content (see Table 7–2) because this impurity was believed to have caused corrosion in cooling jackets of certain thrust chambers. See Ref. 7–14.

Methane (CH₄) is a cryogenic liquid hydrocarbon and the main constituent of ordinary “natural gas.” Liquid methane is abundant and relatively low in cost. Compared to petroleum‐refined hydrocarbons, it has highly reproducible properties. A bipropellant consisting of LOX and methane gives a lower characteristic velocity and/or specific impulse I_s than LOX‐hydrogen but higher than LOX‐kerosene or LOX‐RP‐1, see Table 5–5. Several rocket propulsion organizations in different countries have been developing and ground testing experimental liquid propellant thrust chambers and/or small reaction control thrusters with LOX‐CH₄. Since the density of liquid methane is approximately six times higher than liquid hydrogen, methane fuel tanks can be much smaller and less costly. Methane is being considered as a bipropellant fuel (with LOX) for future rocket engines in large, multistage space launch vehicles for both manned and unmanned missions to Mars. In anticipation of three Mars flights, a new larger oxygen‐methane engine is being developed in the United States; while these investigations are presently ongoing, there have not been any flights (as of 2015, personal communication with J. H. Morehart of The Aerospace Corporation).

Liquid Hydrogen

As shown in Table 5–5, hydrogen when burned with oxygen gives a high performance colorless flame (shock waves may however be visible in the rocket plume). Liquid hydrogen is an excellent regenerative coolant. Of all known fuels, liquid hydrogen is by far the lightest and the coldest, having a specific gravity of 0.07 and a boiling point of about 20 K. This extremely low fuel density requires bulky and large fuel tanks, which necessitate rather large vehicle volumes. The extremely low storage temperatures limit available materials for pumps, cooling jackets, tanks, and piping because many metals become brittle at such temperatures.

Because of their very low temperatures, liquid hydrogen tanks and supply lines have to be well insulated to minimize hydrogen evaporation or the condensation of moisture or air on the outside, with subsequent formations of liquid or solid air and ice. A vacuum jacket has often been used in addition to the insulating materials. All common liquids and gases solidify in liquid hydrogen. Such solid particles in turn plug orifices and valves. Therefore, care must be taken to purge all lines and tanks of air and moisture (flushing with helium and/or creating vacuums) before introducing any liquid hydrogen propellant. Mixtures of liquid hydrogen and solid oxygen or solidified air can actually explode.

Liquid hydrogen is manufactured from gaseous hydrogen by successive compression, cooling, and expansion processes. Hydrogen has two isomers, namely, orthohydrogen and parahydrogen, which differ in the orientation of their nuclear spin state. As hydrogen is liquefied and cooled to lower and lower temperatures, the relative equilibrium composition of ortho‐ and parahydrogen changes. The transformation from orthohydrogen to parahydrogen is exothermic and results in excessive boil‐off, unless complete conversion to parahydrogen is achieved during liquefaction. Parahydrogen is the only stable form of liquid hydrogen.

Hydrogen gas, when mixed with air, is highly flammable and explosive over a wide range of mixture ratios. To avoid this danger, escaping excess hydrogen gas (at tank vent lines) is often intentionally ignited or “flared” in air. Liquid hydrogen is used with liquid oxygen in the Delta IV, Ariane V and ‐HII launch vehicles, Centaur upper stage, and upper stage space engines developed in Japan, Russia, Europe, India, and China.

Hydrogen burning with oxygen forms a nontoxic essentially invisible exhaust. This propellant combination has been applied successfully to space launch vehicles because of its high specific impulse (payload capability usually increases greatly for relatively small increases in specific impulse) even if the low density of liquid hydrogen makes for a very bulky fuel tank, a large vehicle, and relatively higher drag. Studies have shown that, when burned with liquid oxygen, hydrocarbons (such as methane or RP‐1) give a small advantage in space launch vehicle first stages. Here, the higher average propellant density allows a smaller vehicle with less mass and lower drag, which compensates for the lower specific impulse of the hydrocarbon when compared to hydrogen. Therefore, some concepts exist for operating the same booster‐stage rocket engines initially with a hydrocarbon fuel and then switching during flight to hydrogen. Engines using LOX with these two fuels, namely, hydrocarbon and hydrogen, are called tripropellant engines. They have not yet been fully developed or flown. Some work on experimental tripropellant engines was done in Russia, but there is no current effort known to the authors.

Hydrazine (N₂H₄)

Reference 7–15 gives a good discussion of this propellant, which is widely used as a bipropellant fuel as well as a monopropellant. Hydrazine, monomethylhydrazine (MMH), and unsymmetrical dimethylhydrazine (UDMH) have similar physical and thermochemical properties. Hydrazine is a toxic, colorless liquid with a high freezing point (275.16 K or 35.6 °F). Hydrazine tanks, pipes, injectors, catalysts, and valves are usually electrically heated to prevent freezing in cool ground weather or in outer space. Hydrazine has a short ignition delay and is spontaneously ignitable with nitric acid, nitrogen tetroxide, and concentrated hydrogen peroxide.

Pure anhydrous hydrazine is a stable liquid; it has been safely heated to near 416 K. It has been stored in sealed tanks for over 15 years. Space probes with monopropellant hydrazine have been operating for about 40 years while traveling beyond the boundary of the solar system. With impurities or at higher temperatures it decomposes releasing energy. Under pressure shock (blast wave or adiabatic compression) hydrazine vapor or hydrazine mist can decompose at temperatures as low as 367 K. Under some conditions this decomposition can be a violent detonation, and this has caused problems in cooling passages of experimental injectors and thrust chambers. Harmful effects to personnel may result from ingestion, inhalation of its toxic vapors, or prolonged contact with skin. The American Conference of Government Industrial Hygienists (ACGIH) recommended eight‐hour personnel exposure limit is 0.01 ppm or 0.013mg/m³. Hydrazine is a known animal carcinogen and a suspected carcinogen for people.

Hydrazine vapors may form explosive mixtures with air. If hydrazine is spilled on a porous surface or a cloth, spontaneous ignition with air may occur. It reacts with many materials, and care must be exercised to avoid contact with storage materials that cause decomposition (see Refs. 7–15 and 7–16). Tanks, pipes, injectors, catalysts, or valves must be cleaned and free of all traces of impurities. Compatible materials include certain stainless steels (303, 304, 321, or 347), nickel, and the 1100 and 3003 series of aluminum. But iron, copper, and its alloys (such as brass or bronze), monel, magnesium, zinc, and some types of aluminum alloy must be avoided.

Hydrazine cannot be safely used in cooling jackets of bipropellant thrust chambers because it will explode when a certain detonation temperature is exceeded. Such explosions occur when heat transfer is unusually high or immediately after thrust termination when the cooling fluid is stagnant; heat soaking back from hot inner walls can overheat any trapped hydrazine and this may cause a violent decomposition. Presently, bipropellant chambers with hydrazine are radiation cooled (using molybdenum or rhenium metals) and/or are ablatively cooled since internal thrust chamber insulation layers can erode and oxidize. Hydrazine injectors have to be designed without inside stagnant pockets or manifolds where liquid hydrazine can remain just after shutdown. Helium purges may be used to solve this situation. Radiation cooling and ablation cooling are discussed in Chapter 8. Reference 7–2 describes recent work on green‐propellant hydrazine replacements.

Unsymmetrical Dimethylhydrazine [(CH₃)₂NNH₂]

This derivative of hydrazine, abbreviated as UDMH, is often used instead of or in mixtures with hydrazine because it forms a thermally more stable liquid. Furthermore, it has a lower freezing point (215.9 K) and a lower boiling point (335.5 K) than hydrazine itself. When UDMH is burned with an oxidizer it gives only slightly lower values of I_s than pure hydrazine. UDMH has been used when mixed with 30 to 50% hydrazine or with 25% hydrazine hydrate. UDMH is used in many Russian and Chinese small thrusters and some main rocket engines. The ACGIH (Ref. 7–6) recommended eight‐hour personnel exposure limit for vapor is 0.01 ppm, and UDMH is a known animal carcinogen.

Monomethylhydrazine (CH₃NHNH₂)

Monomethylhydrazine, abbreviated as MMH, is being used extensively in U.S. spacecraft rocket engines, particularly in small attitude control engines, usually with N₂O₄ as the oxidizer. It has a better shock resistance to blast waves and a better liquid temperature range than pure hydrazine. Like hydrazine, its vapors are easily ignited in air; flammability limits are from 2.5 to 98% by volume at atmospheric sea‐level pressure and ambient temperature. Any materials compatible with hydrazine are also compatible with MMH. The specific impulse with storable oxidizers usually is 1 or 2% lower with MMH than with N₂H₄.

Both MMH and UDMH are soluble in many hydrocarbons, though pure hydrazine is not. Monomethylhydrazine, when added in quantities of 3 to 15% by volume to hydrazine, has a substantial quenching effect on the explosive decomposition of hydrazine. MMH decomposes at 491 K, whereas pure hydrazine can explode at 369 K when subjected to certain pressure shocks. MMH is a known animal carcinogen and the ACGIH recommended personnel 8‐hour exposure limit is 0.01 ppm. Of all hydrazines, MMH vapor is the most toxic when inhaled.

Hydrazine as a Monopropellant

Hydrazine stores well and is used as monopropellant when decomposed by a suitable solid catalyst; such catalyst often needs to be preheated for fast startups and/or for extending the useful catalyst life. Iridium on a porous alumina base is an effective catalyst at room temperature. At elevated temperature (about 450 K) many other materials decompose hydrazine, including iron, nickel, and cobalt. See Ref. 7–15. Different catalysts and different configurations produce different decomposition products, resulting in gases of varying composition and temperature. Monopropellant hydrazine units are used in gas generators and/or attitude control rockets, and as propellant in electrothermal and arcjet thrusters. A typical hydrazine monopropellant thrust chamber, its injection patterns, and its decomposition reaction are described in Fig. 8–14. Typical operating parameters are shown in Fig. 7–3 and Table 7–1.

The catalytic decomposition of hydrazine can be described ideally as a two‐step process; such a simplified scheme ignores other steps and intermediate products. First, hydrazine (N₂H₄) decomposes into gaseous ammonia (NH₃) and nitrogen (N₂); this reaction is highly exothermic (e.g., it releases heat). Second, hot ammonia decomposes further into nitrogen and hydrogen gases, but this reaction is endothermic and absorbs heat. These simplified reaction steps may be written as

7–4

Here, x is the degree of ammonia dissociation; it is a function of catalyst type, size, and geometry; chamber pressure; and dwell time within the catalyst bed. Figure 7–3 shows several calculated rocket engine parameters for hydrazine monopropellant as a function of x. The shown values are for an ideal thruster at 1000 psia chamber pressure with a nozzle area ratio of 50 expanding at high altitudes. The best specific impulse is attained when little ammonia is allowed to dissociate.

For more than 25 years now a number of investigations have been conducted to find a substitute for hydrazine monopropellant because it is toxic, its freezing point is relatively high (275 K or 2 °C), and it requires heating of all parts it comes in contact with. To date, none of the propellants that have been evaluated so far are fully satisfactory. One possible candidate is a monopropellant (identified as LMP‐1035) made from ammonium dinitramine (ADN, a solid) dissolved in small amounts of water; it was first flown in the European satellite PRISMA, which was launched by the Russian DNEPR launch vehicle on June 15, 2010. This relatively new experimental monopropellant has a higher density than hydrazine, is not toxic, and has an acceptable liquid temperature range but a slightly lower performance. However, LMP‐1035 will decompose only when its catalyst has been preheated to 533 K or 500 °F. This higher decomposition temperature requires extra thermal insulation and this heating may shorten the catalyst life. Another substitute candidate is based on hydroxyl ammonium nitrate (HAN) and this monopropellant has also been under investigation. Both HAN‐ and ADN‐based monopropellants are scheduled for tests in U.S. satellites, but to date neither has been fully proven. Investigations continue on other potential hydrazine monopropellant replacements.

Greek Letters