. The Propellant used by Rockets' Propulsions

. Space Power and Energy Storage

. Why Russian Rocket Engines are so Popular?

. Some gas used in the Rockets

Credit: Kris Holland

Video: A concept for a thermal space solar power satellite by Keith Henson, Steve Nixon and Kris Holland; animated by Kris Holland/Mafic Studios and Anna Nesterova. ©Mafic Studios, Inc., www.maficstudios.com Questions about usage should be directed there.




MIT Science Reporter — "Landing on the Moon" (1966). Credit: From the Vault of MIT

Space Power and Energy Storage

Energy Harvesting - Electrodynamic tether

Many technologies are possible to harvest electric power, in megaWatt, from the space environment for use on CubeSat class vehicles.  That one could be a electrodynamic tethers to scavenged energy from a planetary magnetic field to low temperature thermos-electrics to convert low quality waste heat to electricity.

In the example of the electrodynamic tether, a 20 kilometer (km) tethered satellite system experiment has demonstrated > 2 kW power generation in LEO.

The objective for electrodynamic tether technology is to produce, for a period of years, multi-kW power with a 1 km tether at ~1 amp (A) in the Earth magnetosphere or 5 A in the Jovian magnetosphere. The primary challenge for this is the durability of the tether material.

Electrodynamic tethers can produce energy for orbital spacecraft for missions of indefinite length.

An electrodynamic tether is essentially a long conducting wire extended from a spacecraft. The gravity gradient field pulls the tether taut and tends to orient it along the vertical direction. As it orbits around the Earth, or other planet, it crosses the body’s magnetic field lines at orbital velocity at 7 to 8 km/s. The motion of the conductor across the magnetic field induces a voltage along the length of the tether. This voltage, the motional electromagnetic force (EMF), can be up to several hundred volts per kilometer. Within the ionosphere, free electrons are collected by the tether, producing useful power of up to several kilowatts at the expense of the spacecraft orbital altitude, from which the energy is derived.


Also, to Generate Electrical Power we can use Chemical Hydrogen-Oxygen and Hydrocarbon-Oxygen fuel cells. That is, generally, provide the maximum specific energy for crewed spacecraft (non-propulsion) loads when solar power is unavailable.

Every crewed U.S. spacecraft since the Gemini program has used fuel cells to generate electrical power for vehicle loads.

The last fuel cell to fly in a crewed spacecraft is the alkaline hydrogen-oxygen fuel cell used on the Space Shuttle Orbiter. This alkaline fuel cell was sized to generate 2-12 kWe with reagent-quality hydrogen and oxygen, utilizing active pumps and water separators to manage the product water. These external components limited the life of the fuel cell to about 5,000 hours before requiring refurbishment. The alkaline separator for this fuel cell required asbestos, which is now unavailable, and leakage of the caustic potassium hydroxide (KOH) electrolyte limited the power plant’s life.

When clean hydrogen has been available as a fuel, commercial systems in terrestrial applications now use proton exchange membrane (PEM) technology for systems ranging from < 1 kWe for portable applications to 400 kWe for stationary power.

PEM fuel cells for terrestrial transportation systems range from 1 kWe to 100 kWe. Rather than using corrosive pure oxygen as an oxidant, they use air and rely on gravity and external components to expel product water. The effects of variable hydration and heat on the PEM membrane typically limit the operational life of these plants to ~5,000 hours.

To reduce recurring cost of the platinum catalyst required, in commercial systems, the PEM membranes operate typically at low voltage, i.e., at low efficiency.

Solid oxide fuel cells (SOFCs) are now commercially available for terrestrial applications. Their high temperature operations enable compatibility with fuel reformed from such hydrocarbons as natural gas, but also create challenges for reliable operation, notably with thermal transients during on-off cycling. Commercial systems can also be heavy, while space operations require kilowatts of electrical power generation without a substantial mass penalty.

NASA has three primary applications that require advancements in fuel cell technology:

Electrical power generation from oxygen and hydrogen to power crewed transportation systems including landers and rovers. These systems need to produce nominally 40 kWe at 120 volts (V). Balance-of-plant components (regulators, valves, circulation pumps) are the source of most failure modes in fuel cell power plants of any chemistry, so passive reactant management systems could dramatically improve the reliability of fuel cell systems by eliminating the most common sources of failure. System mass can be reduced by new stack bipolar plate designs and materials, and system efficiency and durability can be improved by new catalyst and membrane materials.

The primary objective for PEM fuel cells is to demonstrate reliable operation for > 10,000 hours with high efficiency (> 75 percent) when operating with propellant-grade hydrogen and oxygen. These fuel cell systems must also operate with reactants stored at nominally 2,000 psi, which requires either a reliable compressor or fuel cell stack that does not leak at elevated pressures.

Fuel cell systems that can directly process residual propellants from landers and fuels generated from ISRU systems will enable the ability to establish sustainable outposts.

Power can effectively be produced from liquid methane-oxygen propulsion storage via high-temperature (e.g., solid oxide) fuel cells, bipropellant turbines or Stirling engines, or a combination thereof. Further, high-temperature SOFCs enable heat rejection systems that are greatly reduced in mass. High-temperature fuel cells (e.g., solid oxide) must reliably operate with systems that are cycled on and off up to 10 times, and therefore must demonstrate durability during thermal cycling and high efficiency operation (> 70 percent) when operating with hydrocarbon fuels such as methane.

Electrical power generation from oxygen and hydrogen or methane carbon monoxide (CO) generated by electrolysis as part of a regenerative fuel cell system is needed for crew transportation systems and surface systems. Such advances are particularly advantageous when integrated with ISRU systems.


Energy Storage Technologies for Future Planetary Science Missions

SOON @ Rockets' Propulsions - Part Five



The Propellants used by Rockets' Propulsions

Liquid Cryogenic

Currently, the most used cryogenic liquid propellants for the in-space transfer stages are the Liquid Oxygen (LO2 or LOX) and Liquid Hydrogen (LH2). However, their storage and transfer can be challenging, in particular, to prevent the boil-off for the longtime missions. 

More about LH2

Consisting of 99.79% of parahydrogen and 0.21% of orthohydrogen, the Liquid Hydrogen (LH2) is the liquid state of the element hydrogen found naturally in the molecular H2 form.

To exist as a liquid, H2 must be cooled below its critical point of 33 Kelvin. However, to be fully liquid without boiling at atmospheric pressure, it needs to be cooled to 20.28 K (−423.17 °F/−252.87 °C).

As in any gas, storing hydrogen as a liquid takes less space than storing it as a gas at normal temperature and pressure. However, its density is very low and, once liquefied, it can be maintained as a liquid in pressurized and thermally insulated containers.

The density of LH2 is only 70.99 g/l, at 20 k, mains just 0.07. Although the specific energy is around twice that of other fuels, that gives it a very low volumetric energy density, many fold lower.

Current liquid rocket fuel, it first cools the nozzle and other parts before being mixed with the oxidizer, usually the Liquid Oxygen (LOX), and burned to produce with traces of ozone and hydrogen peroxide. Almost H2-O2 rocket engines run with fuel-rich gas, the exhaust contains some un-burned hydrogen. It result a reduction of the erosion of the combustion chamber and the nozzle. And, because it reduce the molecular weigh of the exhaust, it can increase the specific impulse.

When mixed with the oxygen, it produce only water vapor. which can be cooled with some LH2. Being harmless to the environment, an engine burning it is "zero emission".

Even with thermally insulated containers, it is difficult to keep such a low temperature, and the hydrogen will gradually leak away, at a rate of about 1% per day. 

Specific impulse: 451 seconds (s). Specific impulse sea level: 391 s. Temperature of Combustion: 2,985 deg Kelvin. Ratio of Specific Heats: 1.26. Characteristic velocity c: 2,435 m/s (7,988 ft/sec). Isp Shifting: 391 sec. Isp Frozen: 388 sec. Mol: 10.00 M (32.00 ft). Oxidizer Density: 1.140 g/cc. Oxidizer Freezing Point: -219 deg Celsius. Oxidizer Boiling Point: -183 deg C. Fuel Density: 0.071 g/cc. Fuel Freezing Point: -259 deg C. Fuel Boiling Point: -253 deg C.

The LOX and the LH2 are used because they provide higher performances for in-space stages thrust needed at about 25,000 pounds-force (lbf). Recently, many works are be done for a larger transfer stages wit with a thrust level up to 50,000.  Already used on the Nuclear thermal rocket, these gas are considered for future Mars missions.

More about LOX

Liquid oxygen (LOX) was the earliest, cheapest, safest, and eventually the preferred oxidizer for large space launchers. The reason is it creates a high specific impulse and it can usually be used with yhe Liquid Hydrogen, the Kerosene or the Methane.

Because it is moderately cryogenic, it can not be suitable for military uses, where the storage of the fuelled missile and a quick launch are required. 

It was used in the very first rocket applications like the V2 missile and Redstone, R-7 Semyorka, Atlas boosters and the ascent stages of the Apollo Saturn rockets. However, modern ICBMs do not use liquid oxygen because its cryogenic properties and the regular need of replenishment to replace boiloff it hard to maintain, as well as to launch quickly.

Liquid nitrogen has a lower boiling point at −196 °C (77 K) than oxygen's −183 °C (90 K), and vessels containing liquid nitrogen can condense oxygen from air: when most of the nitrogen has evaporated from such a vessel there is a risk that liquid oxygen remaining can react violently with organic material. Conversely, liquid nitrogen or liquid air can be oxygen-enriched by letting it stand in open air; atmospheric oxygen dissolves in it, while nitrogen evaporates preferentially.

Liquid oxygen has a pale blue color and is strongly paramagnetic. It can be suspended between the poles of a powerful horseshoe magnet. LOX has a density of 1.141 g/cm3 (1.141 kg/L or 1141 kg/m3) and is cryogenic with a freezing point of 54.36 K (−218.79 °C; −361.82 °F) and a boiling point of 90.19 K (−182.96 °C; −297.33 °F) at 101.325 kPa (760 mmHg). It has an expansion ratio of 1:861 under 1 standard atmosphere (100 kPa) and 20 °C (68 °F). Because of this latter advantage, it is used in some
commercial and military aircraft as transportable source of breathing oxygen.

Due to the difficulty to maintain the Hydrogen as a liquid and its low storage density, some research have been conducted to use the Liquid Methane (LCH4) as an alternative fuel. LCH4 have a boiling point closer to the liquid Oxygen at -263.2° F and a more simple system than the Hydrogen. Also, for missions on Mars or other bodies, Methane can be produced In-Situ and used for reaction control and ascent/descent propulsion systems.

As see, the 1st & 2nd Stages uses LOX/H2 for in-space mission.

SLS Interim Cryogenic Propulsion Stage. Credit: ULA

Blue Origin use Liquid Oxygen and liquefied natural gas (LNG) for its Stages'engines. Credit: Blue Origin

Gelled and metallized fuels are very popular because they improves the performance of rocket systems. Experiments in laboratory has revealed that, they can be an solution for the boil-off problem and the corresponding shift in propellant-loading.

No matter if they don't have yet flown in a space-representative environment, primary candidates have been identified for them. It was found that, the NTO/MMH/Aluminium, the O2/Rocket Propellant(RP)-1(Kerosene-1)/Aluminium, and the Cryogenic O2/H2/Aluminium are the primary targets. Here, the big challenge is gelling the fuels with Aluminium particles, the storage stability and the combustion efficiency. Certainly,  more research is needed to confirm their capabilities.

For now, it was found that, they can increase the rocket Isp and the fuel density. And, because the mix become more viscous, they can reduce the spill radius in an accidental spill and fuel sloshing, as well as reduces the volatility during an accidental low-pressure propellant fires. Finally, they can reduce the leak potential from damaged fuel tanks.

Solid propellants usually comprise pre-mixed oxidizer and fuel. The mix is cast so that the surface area burns at a predetermined and tailored burn rate when ignited. The burn rate generates the thrust and duration required for the mission. For solid propellants, Isp values are normally less than 300 seconds. For space-based solids, hydroxylterminated polybutadiene (HTPB) propellant has been used exclusively in apogee kickmotors and upper stages. Thrust vectoring is controlled by gimballing or gaseous/liquid

Solid rockets are relatively simple to operate and have a high-density Isp. The pre-mixed oxidizer and fuel can be cast into a particular shape to predetermine and tailor burn rates to generate the thrust and duration required for a mission.

The Space Shuttle

A pre-mixed oxidizer and fuel gives normally Solid propellants. When it is the time to start, the mix is throw and the surface area burns at a predetermined and tailored burn rate. This burn generates the thrust and duration in seconds needed for a mission. For solid propellants, Isp values are normally less than 300 seconds.

For space-based solids, as the hydroxyl-terminated polybutadiene (HTPB) propellant, has been used exclusively in apogee kick-motors and upper stages. The thrust vectoring is controlled by gimballing or gaseous/liquid injection.

Some rockets uses a solid fuel and a liquid oxidizer to make their thrusts. These Hybrid rockets have an higher Isp, are safer than solid-propellant rockets, and are less complex and expensive than Liquid rockets. Physically, they are generally bigger than solid rockets because the propellant they uses have a lower density, and so, need more volume.

Hybrid motors have been demonstrated at the 250-kilo pounds-force (klbf) thrust level in ground testing and recent developments in hybrid technology have resulted in significant progress, reducing technology risk, with long-burn-duration firings at the 20 klbf thrust level for an upper stage application.

The Cold and/or the Warm Gas system(S) provides a compact gas propulsion system that can provide small delta-V or small total impulse to a spacecraft system.

Using in space since 1950s, the cold gas flown in small satellites, upper stages, or in human space exploration with extravehicular activities (EVAs).

Switching to a warm gas system will increase the Isp, the mission delta-V and total impulse. There are many options that are flight ready for all categories of micro-propulsion, ranging from about 0.1 Newton for cold gas and 1.07 Newton for hydrazine thrusters to 170 Newton for solid motor options. Micro-hydrazine thrusters are used to produce low thrust levels and minimum-impulse burns for reaction control systems. Solid motors are used to provide precision impulse for deployments, attitude changes, spin up and spin down, and more. Cold or warm gas propulsion systems are used for precise attitude control and precision impulse bits.

Warm gas systems have been used in flight systems for pressurization but not for main propulsion. The principal advantage of using the warm gas version of a cold-gas system is that it requires approximately half the propellant storage tank volume. Gas propulsion systems are typically used for small delta-V rockets or when small total impulse is required. These systems are generally inexpensive and very reliable, and inert gases are inherently non-toxic. Most of the residual risk lies with the high-pressure storage tanks, although good design provides ample margin for safety.

Below, some multi-stages rockets & their propellants used. Credit: FAA



MIT Science Reporter—"Computer for Apollo" (1965). Credit: From the Vault of MIT

Why Russian Rocket Engines are so Popular?

Russian and American rocket engines use liquid oxygen as oxidizer and kerosene or RP-1 (some kind of kerosene) as fuel. So, why Russian Rocket like RD-191 give a specific impulse (Isp) much higher and offering a significant reduction in propellant mass than the US?

Figure: the Russian RD-191 Engine

Because, by the use of a closed cycle system to drive the pump, the Russians obtain a much higher chamber pressure that allows bigger (geometric) expansion ratio and hence a higher exhaust velocity. In that closed cycle system, the (cool) turbine drive gases are fed to the combustion chamber thereby allowing the gases to be used again, but this time to produce thrust.

In contrast, the US designs are based on a gas-generator cycle where an open cycle meaning that the turbine drive gases of the pumps are dumbed separately, without using it to provide thrust (at least not in an optimal way).

Implemented in the 1960's, the Russias were able to use the closed cycle system because at that time they had developed the necessary materials to use oxygen-rich combustion gases from the pre-combustor to drive the turbopump(s).

In the USA NASA part. for long held, they view that the use of
oxygen-rich turbine drive gases induces corrosion problems. And, lately, this view has changed because today we have significant number of materials available that are fully capable of being used as turbine components in hot oxygen-rich gases for service lifetimes of many hours.

Notice that the reason to use oxygen-rich turbine drive gases in a closed cycle LOX-kerosene system is because the kerosene propellant is also used as coolant, with maximum coolant temperatures being limited to about 490 K to prevent varnish deposits.

Integration of the RD-191 engine with an experimental version of the URM-1 booster for the Angara rocket.

Others gas used in the Rockets

The Liquid Oxygen/Kerosene RP-1 (Rocket Propellant-1 or Refined Petroleum-1), or LOX / RP-1 (Kerosene), is a highly refined form of kerosene outwardly similar to jet fuel used as rocket fuel. RP-1 has a lower specific impulse than liquid hydrogen (LH2), but are cheaper, stable at room temperature, far denser, more powerful and has a fraction of the toxicity and carcinogenic hazards of the Hydrazine.

The RP-1, commonly burned with the oxidizer liquid oxygen (LOX), is a fuel used for the first stage boosters of the Soyuz-FG, Zenit, Delta I-III, Atlas, Falcon 9, Antares and Tronador II rockets. It also powered the first stages of the Energia, Titan I, Saturn I and IB, and the historic Saturn V.

Unlike the hydrogen-fueled equivalents, powered boosters are always more compact due to the order-of-magnitude higher density of the propellant. Booster stage configurations are multi-engine and include thrust vector control by nozzle gimballing.

Kerosene is hardly ever used for upper stage in the United States. The existing SOA is manifested instead in multiple Liquide Rocket Engines, most notably the foreign-designed and built heritage RD-180 and the NK-33, both utilizing Oxidizer-Rich Staged Combustion (ORSC) with Isp over 330 seconds.

These foreign-made products are procured by U.S. domestic expendable launch vehicle providers, like Rocketdyne. Domestically, the U.S. industry has only produced the simpler kerosene gas generator cycle engines (a little over 300 seconds) and is thus reliant on foreign countries for higher-performing ORSC versions.

Soviet and Russian rocket-grade kerosene are very similar to RP-1 and are designated T- 1 and RG-1. Densities are higher, 0.82 to 0.85 g/ml, compared to RP-1 at 0.81 g/ml. For a short period, the Soviets achieved even higher densities by super-chilling the kerosene in a rocket’s fuel tanks, but this partially defeated the purpose of using kerosene over other super-chilled fuels. In the case of the Soyuz and other R7-based rockets, the temperature penalty was minor. Facilities were already in place to manage
the vehicle's cryogenic liquid oxygen and liquid nitrogen, both of which are far colder than the kerosene temperature.

The launcher's central kerosene tank is surrounded on four sides and the top by liquid oxygen tanks, the liquid nitrogen tank is nearby at the bottom. The kerosene tanks of the four boosters are relatively small and compact, and also between a liquid oxygen and a liquid nitrogen tank. Thus, once the kerosene was chilled initially, it could remain so for the brief time needed to finish launch preparations.

All told, kerosene engines generate an Isp in the range of 270 to 360 seconds, while hydrogen engines achieve 370 to 465 seconds.

Any hydrocarbon-based fuel when burned produces more air pollution than hydrogen. Hydrocarbon combustion produces carbon dioxide (CO2), toxic carbon monoxide (CO), hydrocarbon (HC) emissions, and oxides of nitrogen (NOx), while hydrogen (H2) reacts with oxygen (O2) to produce only water (H2O), with some un-reacted H2 also released.

An important benefit of a new domestic RP/LOX engine would be reducing or eliminating the reliance on international partners for providing a critical part of a launch propulsion system. Eventually, that will develop a bigger supply of this advanced high performing kerosene engine for many decades within the domestic industrial base.

A large ORSC engine of one million pounds thrust (or higher) is a candidate to replace SLS solid rocket boosters in its 130 metric Tons configuration, and is thus an enabling technology for the human exploration of Mars.

Specific impulse: 353 s. Specific impulse sea level: 300 s. Optimum Oxidizer to Fuel Ratio: 2.56. Temperature of Combustion: 3,670 deg K. Ratio of Specific Heats: 1.24.
Density: 1.02 g/cc. Characteristic velocity c: 1,805 m/s (5,921 ft/sec). Isp Shifting: 301 sec. Isp Frozen: 286 sec. Mol: 23.30 M. Oxidizer Density: 1.140 g/cc. Oxidizer Freezing Point: -219 deg C. Oxidizer Boiling Point: -183 deg C. Fuel Density: 0.806g/cc. Fuel Freezing Point: -73 deg C. Fuel Boiling Point: 147 deg C.
Credit: http://www.astronautix.com/l/loxkerosene.html

The Nitrogen tetroxide/Unsymmetrical dimethyl hydrazine (N2O4/UDMH). The UDMH, a derivative of hydrazine, is often used in hypergolic rocket fuels as a bipropellant in combination with the oxidizer Nitrogen tetroxide (N2O4) and less frequently with Liquid oxygen.

The UDMH has a boiling point of 63º C and N2O4, 21º C, which makes it more storable than cryogenic. So, because it is stable, it can be kept loaded in rocket fuel systems for long periods. Instead UDMH has higher stability than hydrazine, especially at elevated temperatures, it can be used as its replacement or together in a mixture.

UDMH is used in many European, Russian, Indian, and Chinese rocket designs. The Russian Proton, Kosmos-3M, and the Chinese Long March 2F are the most notable users of UDMH
,which is referred to as "heptyl" by Russian engineers. The Titan, GSLV, and Delta rocket families use a mixture of 50% Hydrazine and 50% UDMH, called Aerozine-50, in different stages.

Relative to other storable hypergolics, UDMH/N2O4 has a rather high specific impulse, at about 330 seconds in a vacuum. Other derivative come close, like Aerozine-50 gives 310 seconds in vacuum, but UDMH/N2O4 still the most efficient in terms of specific impulse.

All are very popular propellants used in a variety of launch vehicles, satellites, and on the International Space Station. The combination is especially popular for reaction control thrusters on spacecraft—the ISS's Zvezda Service Module has 34 UDMH thrusters for the orbital station keeping (although the thrusters on visiting spacecraft are usually used instead).

On the space shuttle, a complicated pyrotechnic ignition system ignited the Space Shuttle Main Engine right before liftoff. The chemicals ignite spontaneously when mixed with each other and the flames then ignited the propellants in the combustion chamber.

UDMH/N2O4 is unique in that the combination is hypergolic, which means that, like the pyrotechnic chemicals, it spontaneously ignite when mixed. This removes the need for a separate system.

Spacecrafts and satellites requires hypergolic propellants because their thrusters are used for reaction control, which requires to fire intermittently and for short bursts the thrusters. Due to the complexity in igniting a rocket engine with
pyrotechnics, the ignition sequences is usually measured in multiple seconds, which makesshort burst firings impossible.

Specific impulse: 333 s. Specific impulse sea level: 285 s. Optimum Oxidizer to Fuel Ratio: 2.61. Temperature of Combustion: 3,415 deg K. Ratio of Specific Heats: 1.25. Density: 1.18 g/cc. Characteristic velocity c: 1,720 m/s (5,640 ft/sec). Isp Shifting: 285 sec. Isp Frozen: 273 sec. Oxidizer Density: 1.450 g/cc. Oxidizer Freezing Point: - 11 deg C. Oxidizer Boiling Point: 21 deg C. Fuel Density: 0.793 g/cc. Fuel Freezing Point: -57 deg C. Fuel Boiling Point: 63 deg C. Credit: http://www.astronautix.com/n/n2o4udmh.html

For one, UDMH/N2O4 is extremely toxic. Traditional hydrazine/hypergolics are extremely toxic, corrosive, flammable and needto be handled with extreme care. This complicates ground support equipment and procedures. The combustion of UDMH/N2O4 also produce large quantities of nitrogen oxides, which can further react with water vapor and sulfate in the atmosphere to formsmall particles containing nitric acid.