Galactic Library - User contributions [en]

Photon Rocket

2021-10-04T06:47:49Z

Bruhm:

This is it - the be-all and end-all of reaction drives. No other drive competes with a photon rocket's metrics of output, efficiency, and complexity. Effectively a flashlight of sufficient power to generate practical thrust, the photon rocket is of the same species as chemical and nuclear rockets, but evolved beyond either. Instead of chemical bonds or nuclear fusion and fission, a photon rocket uses one-to-one antimatter-matter annihilation to convert the fuel's mass into radiation. Directed in a single direction, this light is sufficient to create thrust.

==Introduction==

Reaction drives are all about energy and efficiency. The more energy you can extract and apply to a reaction mass, the greater your efficiency, the mightier your rocket. A chemical rocket might combine two compounds, burning them, and using the released energy to eject the fuel's mass as fast as possible. A nuclear rocket extracts that same energy from nuclear fusion or fission, and due to the nature of nuclear reactions, manages a much greater energy density potential. For example, in the case of hydrogen-hydrogen fusion, 0.645% of the reaction's mass is carried away in the form of radiation<ref> https://books.google.com/books?id=Mg4AAAAAMBAJ&pg=PA99 Bulletin of the Atomic Scientists. </ref> - less than a percent of the mass in play becomes light.

A photon rocket converts 100% of its fuel into energy. Reacting 0.5 grams of antimatter with 0.5 grams of matter results in 0 grams of either, and ~9×10^13 Joules (21.5 kilotons-equivalent) of energy. <ref> http://www.projectrho.com/public_html/rocket/antimatterfuel.php Atomic Rockets. </ref> A nuclear bomb's worth of energy is generated with the mass as an average pen cap. Where batteries, chemical fuels and nuclear bombs all store energy in their mass, antimatter is the ultimate battery, because the entirety of its mass is the energy potential in play.

Some designs choose to apply that energy to accelerating matter, ejecting it to create thrust. Instead, a photon rocket uses the generated radiation directly, reflecting it off the rocket's drive and casting into one direction, thrusting the craft the other way. No mass is ejected, just a cone of radiation. It's an elegant, efficient, and incredibly complicated method - after all, we need to reflect that radiation without absorbing almost any of it, lest we become hot, expanding plasma. The radiation in play is of a vast variety, up to and including gamma and neutron radiation that penetrates meters of lead and water, although workable designs strive to generate as little of those as possible. We also need to store antimatter without allowing it to touch matter, for many months or years, in a acceleration-resistant and secure way that doesn't have the crew constantly fearing an instant and unexpected death.

===Theoretical Performance===

The exact mathematics of a photon drive depend on its design. <ref> https://vixra.org/pdf/1201.0026v1.pdf "Matter-Antimatter GeV Gamma Ray Laser Rocket Propulsion" by F. Winterberg </ref> <ref> https://nets2021.ornl.gov/wp-content/uploads/gravity_forms/12-b63a96649a525ab5aa39d607840d9d9f/2021/04/jackson_exoplanet_202104261.pdf "Antimatter-Based Propulsion for Exoplanet Exploration" by Dr. Gerald P. Jackson </ref>

==Design and Function==

===Reflection and Direction===

===Storage===

===Performance / Capabilities / Applications===

==References==

Ion Engine

2021-09-23T01:32:26Z

Bruhm:

Ion engines are an electric thruster type that uses charged particles as propellant. Particle sizes range from subatomic particles to small grains of dust. Propellant is ejected from a nozzle using electrostatic or electromagnetic forces. Generally, they have very high exhaust velocities and little thrust.

Thrust power has to be handled by components such as wires, electrostatic tubes or magnetic coils. These are usually heavy and limited to a low operating temperature, so the specific power of an ion engine is much lower than a thermal rocket. After including the power generating system and the radiators to remove waste heat, we get a propulsion system that accelerates slowly.

The low acceleration of ion engines means spacecraft generally have to use less efficient spiral trajectories near planets with a strong gravitational field, and have to forgo the benefits of Oberth burns. Trying to take off from a planet will usually require an entirely different propulsion system with better thrust. The ion drive’s high exhaust velocity means that an impressive amount of deltaV is possible with a small propellant load. A spaceship with ion engines can accelerate to huge velocities given enough time; the greater the distance required, the likelier it is that an ion drive can get you there in less time.

==Introduction==
Ion engines are not a modern technology, but have only recently been embraced by aerospace companies. Satellite manufacturers are relying on ion engines to perform station-keeping over the course of decades. Interplanetary probes use them for deep space maneuvers, or to cut down on the size of the spacecraft needed to reach planets like Mercury or Jupiter.

This is all thanks to it being a highly scalable solution with a high specific impulse, leading to kilometers per second of deltaV with low mass ratios. The smallest ion engines are the size of your fingertip, and they have no upper limit to how large and powerful they can get.

Most development of ion engines happens at small scales, aimed at CubeSat-sized missions. These engines consume a few watts to a few hundred watts of power, with their thrust being measured in milliNewtons. Their main challenge is efficiency. Converting electrical current from the voltage that is produced by solar panels to the voltage used by the engine, ionizing the propellant flow, energizing the nozzle coils and even resistance in the wires become disproportionately large losses. Less watts come out of the engine as exhaust than is spent doing all these other tasks.

There is also research into making larger ion engines. Existing record-holding designs climb into the hundreds of kilowatts. Their efficiency improves, as does their specific power. However, some designs run into endurance limits. Solid electrodes are eroded after being held at high temperatures, ceramic chambers suffer from sputtering and electric grids wither away when bombarded by ions. Solutions like electrodeless thrusters are necessary to run these engines for long durations.

Ion engines naturally have low thrust due to their high exhaust velocity. The electrical generation system and radiator requirements quickly add mass that reduces the overall specific power of the propulsion system. Low specific power and low thrust means much lower acceleration than a thermal engine of the same power level. Prolonged burns are needed to get anywhere.

Low acceleration limits ion engines are only suited for transfers from one orbit in space to another. Strong gravity bends the trajectories they can take into a series of tight spirals that cost a lot more deltaV than the simple Hohmann transfers a high-thrust engine could use. A spaceship equipped with an ion engine that tries to climb from Low Earth Orbit to reach the Moon could take months and require more than twice the deltaV of a chemical rocket. The low thrust of ion engines prevents them from overcoming the gravity of a planet for a surface takeoff, exploiting the Oberth effect or making use of powered gravitational assists.

Where ion engines shine are missions where long-duration accelerations are key to mission success. DeltaVs in the tens of kilometers per second are possible even with today’s technology. They can drastically reduce the travel time between planets and if given enough time, cover great distances. Their low acceleration is not a major hindrance in accelerating in interplanetary space, overcoming solar gravity,.or around the reduced surface gravities of the numerous small bodies in our solar system.

===Engine Performance===
Example table for Engine Performance for Busek RF Gridded Electrostatic Ion Thrusters
{| class="wikitable"
|+
|Name
|Busek BIT-3
|-
|ISP or Exhaust Velocity
|<22,555 m/s <2300 s
|-
|Thrust
|<1.25 mN
|-
|Mass
|1.28 kg
|-
|Efficiency
|17.6% (Calculated)*
|-
|Drive Power
|<14.09 W (Calculated)**
|-
|Power Consumption
|56-80 W
|-
|T/W
|0.0000996 (Calculated)
|-
|Energy Source (Fuel)
|Electricity
|-
|Propellant
|Iodine
|-
|Reactor
|Grids
|-
|Specific Density
|10.9 W/kg (Calculated)
|}

==Design and Function==
There are three main steps in an ion engine’s operation.
-Delivering propellant
-Creating charged particles
-Accelerating charged particles

Ion engines conduct these steps in a variety of ways, depending on their exact model. The NSTAR ion engine used on the Dawn space probe used xenon gas as propellant. The gas was piped into a chamber where an electron gun knocked off charges from the xenon atoms, giving them a positive charge and making them xenon ions. Once inside the chamber, the ions were held by a magnetic field to improve ionization efficiency. Eventually, the ions would diffuse out of the chamber and into the gap between a pair of charged grids. The voltage difference between the grids accelerated the ions out of the engine, producing thrust, and shunting free electrons back into the chamber. A second electron gun fired electrons into the exhaust to neutralize it.

While NSTAR’s xenon engine is efficient, there are several other ways to achieve the three steps listed above. For example, a liquid propellant could be sprayed in the form of droplets, charged particles could be created by the action of radiofrequency waves and the propellant could be accelerated by powerful magnetic pulses. Each method has its own advantages and disadvantages, but in general, designers of ion engines strive to minimize the portion of electrical energy spent ionizing the propellant and maximizing how much is spent sending it out the nozzle. This has an impact on the choice of propellant. Ion engines that use mercury need 5.02 MJ/kg, while xenon requires 8.91 MJ/kg. A thruster that aims for a higher exhaust velocity at the expense of thrust has a lower propellant flow; this allows mission planners to spend more of their launch payload on useful hardware, rather than propellant.

Here is a list of elements by their ionization energy in MJ/kg:

Note that ion engines can be designed to use a huge number of propellants, including molecules, dust particles, liquid metals and more. Ionization energy is an important parameter but many other factors come into play.

For example, many of the elements with the lowest ionization energy per kilogram are highly radioactive and solid up to extreme temperatures. They would be a pain to handle! Other options may be toxic (mercury) or carcinogenic (cadmium). Erosion from reactive species like potassium must be considered. As is the availability and ease of production of a propellant if ISRU is something you want. Xenon may be a great choice for most electric thrusters, but it is much harder to find in our solar system than nitrogen.

Some ion engine designs are flexible enough to use multiple propellants, or mixes that meet certain specifications. For example, airbreathing electric ramjets use different nitrogen, oxygen and hydrogen ratios depending on which orbital altitude they operate at.

With all these options and possibilities in mind, we can reduce ion engines into two main categories:
-Electrostatic thrusters
-Electromagnetic thrusters

A third, less explored category is the radioisotope ion thruster.

==Engine Design==
The categories differ in how the ions are accelerated.

===Electrostatic Ion Thruster:===
Electrostatic thrusters accelerate ions via large voltage differences. Charged particles feel an attraction to a charge of the opposite sign and a repulsion to a charge of the same sign, resulting in a force called the Coulomb force.

An electron with a charge of -1 leaving a hot cathode to an anode with a 100 V voltage difference between them will acquire an energy of 100 electronVolts (or eV). A nanogram sized speck of dust with a million positive charges moving up a 10 V gradient will acquire an energy of 10 million electronVolts (or MeV).

The equation for how much energy E in eV a charged particle gains is:
 E = qV
 With E the energy in joules, q is the particle’s charge in coulombs, and V the voltage in Volts.

The electron is a lightweight subatomic particle. 100 eV is enough to get it moving at 5929 km/s. The dust particle is relatively heavy. 10 MeV will just push it to 1.79 m/s.

These figures are derived from the equation for kinetic energy.
 E = 0.5mv^2
 Or rearranged
 v = (2E/m)^0.5
 With E the particle’s energy in joules, m the particle’s mass in kg, and v the velocity in m/s.

There are four types of electrostatic thruster that use this principle:

*The gridded electrostatic ion thruster [link], which uses electrostatic grids to accelerate the propellant. Charged particles are subject to an electric force when travelling between the grids but not outside the grids.

*The electrostatic colloid thruster [link], which uses droplets of ionic liquid produced by an electrospray ionization process. An electrode provides the electric field to extract the droplets and accelerate them.

*The Hall effect ion thruster [link], which uses a voltage between a cylindrical anode and a negatively charged plasma cathode. Propellant is introduced and ionized at the anode, and accelerates after reaching the cathode, picking up electrons and becoming a neutral gas so that it is not pulled back into the engine.

*The field emission electric propulsion thruster [link], which extracts ions from liquid metal using very high electric fields. Liquid metals like caesium or iridium flow into the tip of a fountain-pen-like structure to make this process easier.

There is a form of electrostatic colloid thruster that uses the beta emissions of a radioactive substance to generate the voltages necessary to eject droplets of ionic liquid.

===Electromagnetic Ion Thruster:===
There are even more ways to accelerate ions using magnetic fields instead of electric fields. Charged particles are sensitive to the Lorentz force, the same force that ejects projectiles out of a railgun. Other electromagnetic ion thrusters use the ponderomotive force, the result of oscillating magnetic fields, or have found some clever way to use magnetic fields.

Magnetic fields push on negatively and positively charged particles equally. This means they put on both ions and electrons, which together form a plasma, and is the reason this type of electric thruster is also called a plasma thrusters.

There are 8 main types of electromagnetic ion thrusters in summary:

*The pulsed plasma thruster [link], which uses an electric arc to vaporize propellant (usually a block of teflon or PTFE) into plasma. Due to the force of the vaporization, the propellant plasma is ejected into the gap between an electrode and anode, where an electrical current runs through it and causes it to be further accelerated by the Lorentz force.

*The ELF pulsed plasmoid thruster [link], which generates a plasmoid (a smoke torus of plasma) from propellant and expands it down an electrically conducting nozzle, with directed kinetic energy (KE) generated from image currents in the nozzle due to interactions with the plasmoid. It is also electrodeless.
**The alfvenic pulsed plasmoid thruster [link], which uses magnetic reconnection (the same kind that causes solar flares) to spontaneously generate low-temperature plasmoids. The outgoing plasmoid is an alfvenic outflow from the reconnection site, and so depends on the magnetic field strength rather than the propellant ion species mass for thrust.

*The pulsed inductive thruster [link], which uses pulses instead of continuous thrust and can run on high power levels (on the order of megawatts). Using a coil and conical propellant emitter, large built up charges are released behind the coil. This creates an electric current, and induces a magnetic field in the propellant. Together, the electric and magnetic fields accelerate the propellant through the Lorentz force.

*The magnetoplasmadynamic thruster [link], which ionizes the propellant via the electric field between an anode and a cathode. The plasma conducts electricity between the anode and cathode, closing the circuit, and creating a new magnetic field, which with the electric field, accelerates the propellant through the Lorentz force.

*The electrodeless plasma thruster [link], which does not use anode or cathode electrodes. Propellant is ionized through electromagnetic waves, and oscillating electric and magnetic fields in another chamber accelerate the propellant due to the ponderomotive force. The separation of ionization and acceleration steps allows for throttling.

*The helicon double-layer thruster [link], which uses radio waves emitted from an antenna to ionize the propellant into a plasma and excite a helicon wave (a low frequency EM wave bounded inside the plasma), further heating it. A special kind of magnetic nozzle accelerates the propellant.

*The VASIMR [link], which functions similarly to the helicon double-layer thruster but uses an additional second step to heat the plasma to 1 million K through ion cyclotron resonance heating. It uses superconductors to magnetically confine the plasma, and thrust is generated when the plasma escapes confinement as the exhaust. VASIMR has mechanisms that allow it to trade lower exhaust velocity and Isp for higher thrust.

==History / Development==
[Link to specific history article for Ion Engines]

'''[This section should be in its own specific history article, so a link would be provided here instead of the text]'''

The conceptual history of ion engines dates back to 1911, by Konstantin Tsiolkovsky, the first person to publish a paper on the subject. He recommended it for near-vacuum conditions at altitude, however thrust was also seen with ionized air streams at atmospheric pressure.

Hermann Oberth later reignited conversation with his “Ways to Spaceflight” book in 1923, talking about its advantages in mass ratios and mass savings for payloads. He predicted its usage in spacecraft propulsion and attitude control, as well advocated for the electrostatic kind of ion thruster.

By 1959, a gridded electrostatic ion thruster had been built by Harold R. Kaufman at NASA Glenn Research Facilities, using mercury as its propellant. Suborbital tests were conducted throughout the 1960s, and an orbital test was done in 1970, the SERT-1 and SERT-2 demonstration missions.

The Hall effect thruster design, studied by both the United States and the Soviet Union, had been used by the Soviets from 1972 to late 1990s for attitude control. It is estimated that 100-200 engines were used on Soviet and Russian missions. By 1992, the designs had been brought to the West, thanks to a team of electric propulsion specialists from the Ballistic Missile Defense Organization (BMDO) visiting Soviet labs.

==Worked Examples==
===Electrostatic Ion Thruster:===
Gridded Electrostatic Ion Thruster -
NSTAR, NeXT

Colloid Thruster -
Busek Electrospray Thruster Bookmark

Hall Effect Ion Thruster -
PPS-1350, on the SMART-1 mission. AEPS

Field Emission Electric Propulsion Thruster -
IFM Nano

===Electromagnetic Ion Thruster / Plasma Propulsion Engine:===
Pulsed Inductive Thruster -
Several laboratory examples built by NGST.

Magnetoplasmadynamic Thruster -
MPD thrusters flown on EPEX (Japan) mission

Helicon Double-Layer Thruster -
Prototype built by Dr Christine Charles

VASIMR -
VX-10, VX-50, VX-100, VX-200. Built by Ad Astra. (not flown)

==Additional Reading==
Links from ion engine design section
Links from history section

https://cdn.discordapp.com/attachments/886253956777533450/887757110501318757/POTENTIALITIES_OF_THE_RADIOISQTOPE_ELECTROSTATIC_PROPULSION_SYSTEM.pdf
Further reading for the radioisotope ion thruster.

==Additional References==
https://satsearch.co/products/busek-bit-3
 http://www.projectrho.com/public_html/rocket/enginelist.php
 https://www.busek.com/bet-300p

https://en.wikipedia.org/wiki/Ion_thruster
 https://en.wikipedia.org/wiki/NASA_Solar_Technology_Application_Readiness
 https://www1.grc.nasa.gov/space/sep/gridded-ion-thrusters-next-c/

https://en.wikipedia.org/wiki/Gridded_ion_thruster
 https://en.wikipedia.org/wiki/Hall-effect_thruster
 https://en.wikipedia.org/wiki/Field-emission_electric_propulsion

https://en.wikipedia.org/wiki/Pulsed_inductive_thruster
 https://en.wikipedia.org/wiki/Magnetoplasmadynamic_thruster
 https://en.wikipedia.org/wiki/Electrodeless_plasma_thruster
 https://en.wikipedia.org/wiki/Helicon_double-layer_thruster
 https://en.wikipedia.org/wiki/Variable_Specific_Impulse_Magnetoplasma_Rocket
 https://en.wikipedia.org/wiki/Microwave_electrothermal_thruster

https://en.wikipedia.org/wiki/SERT-1
 https://spacemath.gsfc.nasa.gov/weekly/5Page64.pdf
 Electrostatic Ion Engine Math

==Credit:==
To Tshhmon for writing the article
 To SOPHONT SIMP for test-reading
 To KRKIIIIII for test-reading
 To AdAstraGames for test-reading and some editing

Hydrolox Engine

2021-09-23T01:10:23Z

Bruhm:

[[Image:Common Extensible Cryogenic Engine.jpg|thumb|RL10]]

Of all the fuel and oxidizer combinations that do not actively try to eat their way out of propellant tanks, hydrogen and oxygen have the highest performance. The chemical reaction of the two propellants produces a very hot flame and the exhaust, water, exits at a higher velocity than practically anything else. In fact, it burns so hot that most engines move away from the ideal ratio of hydrogen to oxygen to reduce the temperatures inside their combustion chambers.

For rockets that want to be as lightweight as possible, this is the propellant mix to use. But they won’t be the smallest possible as hydrogen is extremely un-dense and deeply cryogenic. This makes designing a propellant tank for it difficult enough, if the hydrogen didn’t also try to leak out from between the atoms of the tank walls.

==Introduction==
Since the early 1960s, hydrogen and oxygen rockets have been the propulsion system with the highest exhaust velocity to anyone designing a launch vehicle. Trying to get high exhaust velocity meant sacrificing all useful thrust, as with an electric thruster, using much scarier chemicals, like fluorine, or developing a very different propulsion system such as nuclear rockets.

Hydrogen-oxygen or ‘hydrolox’ has created the highest performance stages we have ever launched, such as Saturn V’s S-IVB third stage or the Centaur V upper stage.

From a simplistic standpoint, there is no reason to use anything but hydrolox mixes. The real world is not so simple. Other factors come into play that reduce the attractiveness of hydrolox engines. They can be critical obstacles, like the huge volume of propellant tanks meant to contain liquid hydrogen with a density of just 70 kg/m^3 when liquid at near boiling point creating too much drag for spaceplanes to accelerate, to minor things like liquid hydrogen being more expensive than kerosene.

Many missions won’t even make use of hydrolox engines’ superior exhaust velocity. Low deltaV requirements means that performance is wasted. Long duration missions are also less suited to hydrolox propulsion, as storing liquid hydrogen against the heat of unfiltered sunlight beating down on gigantic propellant tanks is a real pain.

===Engine Performance===
This is an example based on RS-25D engine running at its maximal 109% throttle rating.
{| class="wikitable"
|+
|Name
|RS-25D
|-
|ISP or Exhaust Velocity
|366s <-> 3590 m/s (sea level) 452s <-> 4434 m/s (vacuum)
|-
|Thrust
|1,860,000 N (sea level) 2,279,000 N (vacuum)
|-
|Efficiency
|71%
|-
|Drive Power
|5052 MW
|-
|T/W
|73.1 (vacuum)
|-
|Energy Source (Fuel)
|Hydrogen + Oxygen at 6:1 ratio
|-
|Propellant
|Water exhaust
|-
|Reactor
|Combustion chamber
|-
|Power Density
|1590 kW/kg
|}

==Design and Function==
Hydrogen-oxygen engines are, like most chemical engines, fairly simple. You get some hydrogen, you get some oxygen, you pump them into a chamber, you introduce a spark or other ignition source, and steam comes out the back glowing with a very pretty clear flame, sometimes with varying minor coloration.

Hydrolox engines are almost universally fuel-rich and make great use of the excess hydrogen by using it as a coolant for regenerative cooling. This is when propellant flow is used to extract heat from the hottest parts of the engine, such as the nozzle throat, and then adding it into the reaction chamber to improve overall efficiency.

The biggest source of design variations is in how the propellant pumps are powered.

The J-2 used a gas generator, having a secondary burner to power the turbopumps. The J-2S cut that, by tapping off a portion of the main chamber gas flow, and avoiding a separate combustion chamber. You can get more complex, like the multi-stage booster pumps the SSME used to get to nearly 207 bars of pressure. Alternately, you can go dead simple--the expander cycle cleverly manages to harvest all the pumping energy it needs for low-pressure vacuum engine operation from the hydrogen used to cool the chamber and nozzle bell, meaning the turbine inlet gas is only mildly warm. That’s perfect if you want to use turbopump blades made out of a lightweight and strong material that can’t handle high temperature gases, like aluminium alloys.

==Performance / Capabilities / Applications==
The main benefit of hydrolox is its performance: even a moderate pressure engine with a vacuum nozzle can break 440 seconds of specific impulse. No other engine type in present-day service can reach those levels of performance. At the same time, hydrogen is a fantastic coolant, more than an order of magnitude better than any other fuel, meaning it’s possible to build some incredibly clever engine cycles without problems like turbine temperature or propellant “coking” of hydrocarbons.

As an example of the cleverness this allows, the turbine inlet temperature on the expander cycle RL-10 is below the boiling point of water, more like a kitchen appliance than the complex high-temperature turbomachinery needed in other fuels.

Similar endurance and robustness was seen with the J-2 engines for Saturn. The ''Space Shuttle Decision'' by T A Heppenheirner quotes Rocketdyne's Paul Castenholz, one of the engineers managing J-2 development. "We never wore out an engine of the J-2 type. We could run it repeatedly; there was no erosion of the chamber, no damage to the turbine blades. If you looked at a J-2 after a hot firing, you would not see any difference from before that firing. The injectors always looked new; there was no erosion or corrosion on the injectors.” As Heppenheimer reports, this was inspired by examples like a single test engine which ran for 103 starts and 6.5 hours without overhaul. Higher performance hydrolox engines like the Space Shuttle Main Engine (SSME) have a more complex reputation, though more recent tests of the AR-22 variant demonstrated 10 firings in 10 days with no major overhaul. With good performance from even relatively simple engines, it’s no wonder that rocket pioneers like Werner Von Braun were swayed to the Cult of Hydrolox from initial skepticism. For more than half a century, it’s been the propellant of choice for upper and sustainer stages for most American launch vehicles from Atlas to Delta to the Space Shuttle. For that reason, if you want performance with near-term engines, hydrolox has a lot to recommend it.

The downside of hydrolox is the blasted hydrogen. Its performance comes from hydrogen’s low molecular mass--about the smallest anything can get and be legitimately called a molecule--but that also means it has to be incredibly cold to be liquid, at about 20 C colder than liquid oxygen. That’s troublesome for long duration storage, making “boiloff” the word of the day. Heat will get to your hydrogen any way it can, and before long, you won’t have any left. Even once kept liquid, the density is incredibly poor. That means heavier and physically larger tanks for the same performance. Even with hydrogen’s improved specific impulse, the Hydrolox Delta IV lower stage had to be five meters in diameter to lift the same payloads as the 3.8m Atlas V, since hydrolox has about a third the mixed density of kerolox.

==Worked Example(s)==
Planned to go here:

*RL-10 expander
*J-2S gas-gen
*SSME
*Maybe M-1 massive gas gen?

==Additional References==
GRAB SHUTTLE DECISION LINKS

Hydrolox Engine

2021-09-23T01:09:33Z

Bruhm:

[[Image:Common Extensible Cryogenic Engine.jpg|thumb|RL10]]

Of all the fuel and oxidizer combinations that do not actively try to eat their way out of propellant tanks, hydrogen and oxygen have the highest performance. The chemical reaction of the two propellants produces a very hot flame and the exhaust, water, exits at a higher velocity than practically anything else. In fact, it burns so hot that most engines move away from the ideal ratio of hydrogen to oxygen to reduce the temperatures inside their combustion chambers.

For rockets that want to be as lightweight as possible, this is the propellant mix to use. But they won’t be the smallest possible as hydrogen is extremely un-dense and deeply cryogenic. This makes designing a propellant tank for it difficult enough, if the hydrogen didn’t also try to leak out from between the atoms of the tank walls.

==Introduction==
Since the early 1960s, hydrogen and oxygen rockets have been the propulsion system with the highest exhaust velocity to anyone designing a launch vehicle. Trying to get high exhaust velocity meant sacrificing all useful thrust, as with an electric thruster, using much scarier chemicals, like fluorine, or developing a very different propulsion system such as nuclear rockets.

Hydrogen-oxygen or ‘hydrolox’ has created the highest performance stages we have ever launched, such as Saturn V’s S-IVB third stage or the Centaur V upper stage.

From a simplistic standpoint, there is no reason to use anything but hydrolox mixes. The real world is not so simple. Other factors come into play that reduce the attractiveness of hydrolox engines. They can be critical obstacles, like the huge volume of propellant tanks meant to contain liquid hydrogen with a density of just 70 kg/m^3 when liquid at near boiling point creating too much drag for spaceplanes to accelerate, to minor things like liquid hydrogen being more expensive than kerosene.

Many missions won’t even make use of hydrolox engines’ superior exhaust velocity. Low deltaV requirements means that performance is wasted. Long duration missions are also less suited to hydrolox propulsion, as storing liquid hydrogen against the heat of unfiltered sunlight beating down on gigantic propellant tanks is a real pain.

===Engine Performance===
This is an example based on RS-25D engine running at its maximal 109% throttle rating.
{| class="wikitable"
|+
|Name
|RS-25D
|-
|ISP or Exhaust Velocity
|366s <-> 3590 m/s (sea level) 452s <-> 4434 m/s (vacuum)
|-
|Thrust
|1,860,000 N (sea level) 2,279,000 N (vacuum)
|-
|Efficiency
|71%
|-
|Drive Power
|5052 MW
|-
|T/W
|73.1 (vacuum)
|-
|Energy Source (Fuel)
|Hydrogen + Oxygen at 6:1 ratio
|-
|Propellant
|Water exhaust
|-
|Reactor
|Combustion chamber
|-
|Power Density
|1590 kW/kg
|}

==Design and Function==
Hydrogen-oxygen engines are, like most chemical engines, fairly simple. You get some hydrogen, you get some oxygen, you pump them into a chamber, you introduce a spark or other ignition source, and steam comes out the back glowing with a very pretty clear flame, sometimes with varying minor coloration.

Hydrolox engines are almost universally fuel-rich and make great use of the excess hydrogen by using it as a coolant for regenerative cooling. This is when propellant flow is used to extract heat from the hottest parts of the engine, such as the nozzle throat, and then adding it into the reaction chamber to improve overall efficiency.

The biggest source of design variations is in how the propellant pumps are powered.

The J-2 used a gas generator, having a secondary burner to power the turbopumps. The J-2S cut that, by tapping off a portion of the main chamber gas flow, and avoiding a separate combustion chamber. You can get more complex, like the multi-stage booster pumps the SSME used to get to nearly 207 bars of pressure. Alternately, you can go dead simple--the expander cycle cleverly manages to harvest all the pumping energy it needs for low-pressure vacuum engine operation from the hydrogen used to cool the chamber and nozzle bell, meaning the turbine inlet gas is only mildly warm. That’s perfect if you want to use turbopump blades made out of a lightweight and strong material that can’t handle high temperature gases, like aluminium alloys.

==Performance / Capabilities / Applications==
The main benefit of hydrolox is its performance: even a moderate pressure engine with a vacuum nozzle can break 440 seconds of specific impulse. No other engine type in present-day service can reach those levels of performance. At the same time, hydrogen is a fantastic coolant, more than an order of magnitude better than any other fuel, meaning it’s possible to build some incredibly clever engine cycles without problems like turbine temperature or propellant “coking” of hydrocarbons.

As an example of the cleverness this allows, the turbine inlet temperature on the expander cycle RL-10 is below the boiling point of water, more like a kitchen appliance than the complex high-temperature turbomachinery needed in other fuels.

Similar endurance and robustness was seen with the J-2 engines for Saturn. The [i]Space Shuttle Decision[/i] by T A Heppenheirner quotes Rocketdyne's Paul Castenholz, one of the engineers managing J-2 development. "We never wore out an engine of the J-2 type. We could run it repeatedly; there was no erosion of the chamber, no damage to the turbine blades. If you looked at a J-2 after a hot firing, you would not see any difference from before that firing. The injectors always looked new; there was no erosion or corrosion on the injectors.” As Heppenheimer reports, this was inspired by examples like a single test engine which ran for 103 starts and 6.5 hours without overhaul. Higher performance hydrolox engines like the Space Shuttle Main Engine (SSME) have a more complex reputation, though more recent tests of the AR-22 variant demonstrated 10 firings in 10 days with no major overhaul. With good performance from even relatively simple engines, it’s no wonder that rocket pioneers like Werner Von Braun were swayed to the Cult of Hydrolox from initial skepticism. For more than half a century, it’s been the propellant of choice for upper and sustainer stages for most American launch vehicles from Atlas to Delta to the Space Shuttle. For that reason, if you want performance with near-term engines, hydrolox has a lot to recommend it.

The downside of hydrolox is the blasted hydrogen. Its performance comes from hydrogen’s low molecular mass--about the smallest anything can get and be legitimately called a molecule--but that also means it has to be incredibly cold to be liquid, at about 20 C colder than liquid oxygen. That’s troublesome for long duration storage, making “boiloff” the word of the day. Heat will get to your hydrogen any way it can, and before long, you won’t have any left. Even once kept liquid, the density is incredibly poor. That means heavier and physically larger tanks for the same performance. Even with hydrogen’s improved specific impulse, the Hydrolox Delta IV lower stage had to be five meters in diameter to lift the same payloads as the 3.8m Atlas V, since hydrolox has about a third the mixed density of kerolox.

==Worked Example(s)==
Planned to go here:

*RL-10 expander
*J-2S gas-gen
*SSME
*Maybe M-1 massive gas gen?

==Additional References==
GRAB SHUTTLE DECISION LINKS

Hydrolox Engine

2021-09-23T01:08:59Z

Bruhm:

[[Image:Common Extensible Cryogenic Engine.jpg|thumb|RL10]]

Of all the fuel and oxidizer combinations that do not actively try to eat their way out of propellant tanks, hydrogen and oxygen have the highest performance. The chemical reaction of the two propellants produces a very hot flame and the exhaust, water, exits at a higher velocity than practically anything else. In fact, it burns so hot that most engines move away from the ideal ratio of hydrogen to oxygen to reduce the temperatures inside their combustion chambers.

For rockets that want to be as lightweight as possible, this is the propellant mix to use. But they won’t be the smallest possible as hydrogen is extremely un-dense and deeply cryogenic. This makes designing a propellant tank for it difficult enough, if the hydrogen didn’t also try to leak out from between the atoms of the tank walls.

==Introduction==
Since the early 1960s, hydrogen and oxygen rockets have been the propulsion system with the highest exhaust velocity to anyone designing a launch vehicle. Trying to get high exhaust velocity meant sacrificing all useful thrust, as with an electric thruster, using much scarier chemicals, like fluorine, or developing a very different propulsion system such as nuclear rockets.

Hydrogen-oxygen or ‘hydrolox’ has created the highest performance stages we have ever launched, such as Saturn V’s S-IVB third stage or the Centaur V upper stage.

From a simplistic standpoint, there is no reason to use anything but hydrolox mixes. The real world is not so simple. Other factors come into play that reduce the attractiveness of hydrolox engines. They can be critical obstacles, like the huge volume of propellant tanks meant to contain liquid hydrogen with a density of just 70 kg/m^3 when liquid at near boiling point creating too much drag for spaceplanes to accelerate, to minor things like liquid hydrogen being more expensive than kerosene.

Many missions won’t even make use of hydrolox engines’ superior exhaust velocity. Low deltaV requirements means that performance is wasted. Long duration missions are also less suited to hydrolox propulsion, as storing liquid hydrogen against the heat of unfiltered sunlight beating down on gigantic propellant tanks is a real pain.

===Engine Performance===
This is an example based on RS-25D engine running at its maximal 109% throttle rating.
{| class="wikitable"
|+Engine Performance
|-
|Name
|RS-25D
|-
|ISP or Exhaust Velocity
|366s <-> 3590 m/s (sea level) 452s <-> 4434 m/s (vacuum)
|-
|Thrust
|1,860,000 N (sea level) 2,279,000 N (vacuum)
|-
|Efficiency
|71%
|-
|Drive Power
|5052 MW
|-
|T/W
|73.1 (vacuum)
|-
|Energy Source (Fuel)
|Hydrogen + Oxygen at 6:1 ratio
|-
|Propellant
|Water exhaust
|-
|Reactor
|Combustion chamber
|-
|Power Density
|1590 kW/kg
|}

==Design and Function==
Hydrogen-oxygen engines are, like most chemical engines, fairly simple. You get some hydrogen, you get some oxygen, you pump them into a chamber, you introduce a spark or other ignition source, and steam comes out the back glowing with a very pretty clear flame, sometimes with varying minor coloration.

Hydrolox engines are almost universally fuel-rich and make great use of the excess hydrogen by using it as a coolant for regenerative cooling. This is when propellant flow is used to extract heat from the hottest parts of the engine, such as the nozzle throat, and then adding it into the reaction chamber to improve overall efficiency.

The biggest source of design variations is in how the propellant pumps are powered.

The J-2 used a gas generator, having a secondary burner to power the turbopumps. The J-2S cut that, by tapping off a portion of the main chamber gas flow, and avoiding a separate combustion chamber. You can get more complex, like the multi-stage booster pumps the SSME used to get to nearly 207 bars of pressure. Alternately, you can go dead simple--the expander cycle cleverly manages to harvest all the pumping energy it needs for low-pressure vacuum engine operation from the hydrogen used to cool the chamber and nozzle bell, meaning the turbine inlet gas is only mildly warm. That’s perfect if you want to use turbopump blades made out of a lightweight and strong material that can’t handle high temperature gases, like aluminium alloys.

==Performance / Capabilities / Applications==
The main benefit of hydrolox is its performance: even a moderate pressure engine with a vacuum nozzle can break 440 seconds of specific impulse. No other engine type in present-day service can reach those levels of performance. At the same time, hydrogen is a fantastic coolant, more than an order of magnitude better than any other fuel, meaning it’s possible to build some incredibly clever engine cycles without problems like turbine temperature or propellant “coking” of hydrocarbons.

As an example of the cleverness this allows, the turbine inlet temperature on the expander cycle RL-10 is below the boiling point of water, more like a kitchen appliance than the complex high-temperature turbomachinery needed in other fuels.

Similar endurance and robustness was seen with the J-2 engines for Saturn. The [i]Space Shuttle Decision[/i] by T A Heppenheirner quotes Rocketdyne's Paul Castenholz, one of the engineers managing J-2 development. "We never wore out an engine of the J-2 type. We could run it repeatedly; there was no erosion of the chamber, no damage to the turbine blades. If you looked at a J-2 after a hot firing, you would not see any difference from before that firing. The injectors always looked new; there was no erosion or corrosion on the injectors.” As Heppenheimer reports, this was inspired by examples like a single test engine which ran for 103 starts and 6.5 hours without overhaul. Higher performance hydrolox engines like the Space Shuttle Main Engine (SSME) have a more complex reputation, though more recent tests of the AR-22 variant demonstrated 10 firings in 10 days with no major overhaul. With good performance from even relatively simple engines, it’s no wonder that rocket pioneers like Werner Von Braun were swayed to the Cult of Hydrolox from initial skepticism. For more than half a century, it’s been the propellant of choice for upper and sustainer stages for most American launch vehicles from Atlas to Delta to the Space Shuttle. For that reason, if you want performance with near-term engines, hydrolox has a lot to recommend it.

The downside of hydrolox is the blasted hydrogen. Its performance comes from hydrogen’s low molecular mass--about the smallest anything can get and be legitimately called a molecule--but that also means it has to be incredibly cold to be liquid, at about 20 C colder than liquid oxygen. That’s troublesome for long duration storage, making “boiloff” the word of the day. Heat will get to your hydrogen any way it can, and before long, you won’t have any left. Even once kept liquid, the density is incredibly poor. That means heavier and physically larger tanks for the same performance. Even with hydrogen’s improved specific impulse, the Hydrolox Delta IV lower stage had to be five meters in diameter to lift the same payloads as the 3.8m Atlas V, since hydrolox has about a third the mixed density of kerolox.

==Worked Example(s)==
Planned to go here:

*RL-10 expander
*J-2S gas-gen
*SSME
*Maybe M-1 massive gas gen?

==Additional References==
GRAB SHUTTLE DECISION LINKS

Hydrolox Engine

2021-09-23T01:06:56Z

Bruhm:

[[Image:Common Extensible Cryogenic Engine.jpg|thumb|RL10]]

Of all the fuel and oxidizer combinations that do not actively try to eat their way out of propellant tanks, hydrogen and oxygen have the highest performance. The chemical reaction of the two propellants produces a very hot flame and the exhaust, water, exits at a higher velocity than practically anything else. In fact, it burns so hot that most engines move away from the ideal ratio of hydrogen to oxygen to reduce the temperatures inside their combustion chambers.

For rockets that want to be as lightweight as possible, this is the propellant mix to use. But they won’t be the smallest possible as hydrogen is extremely un-dense and deeply cryogenic. This makes designing a propellant tank for it difficult enough, if the hydrogen didn’t also try to leak out from between the atoms of the tank walls.

{| class="wikitable floatright"
|+Engine Performance
|-
|Name
|RS-25D
|-
|ISP or Exhaust Velocity
|366s <-> 3590 m/s (sea level) 452s <-> 4434 m/s (vacuum)
|-
|Thrust
|1,860,000 N (sea level) 2,279,000 N (vacuum)
|-
|Efficiency
|71%
|-
|Drive Power
|5052 MW
|-
|T/W
|73.1 (vacuum)
|-
|Energy Source (Fuel)
|Hydrogen + Oxygen at 6:1 ratio
|-
|Propellant
|Water exhaust
|-
|Reactor
|Combustion chamber
|-
|Power Density
|1590 kW/kg
|}

==Introduction==
Since the early 1960s, hydrogen and oxygen rockets have been the propulsion system with the highest exhaust velocity to anyone designing a launch vehicle. Trying to get high exhaust velocity meant sacrificing all useful thrust, as with an electric thruster, using much scarier chemicals, like fluorine, or developing a very different propulsion system such as nuclear rockets.

Hydrogen-oxygen or ‘hydrolox’ has created the highest performance stages we have ever launched, such as Saturn V’s S-IVB third stage or the Centaur V upper stage.

From a simplistic standpoint, there is no reason to use anything but hydrolox mixes. The real world is not so simple. Other factors come into play that reduce the attractiveness of hydrolox engines. They can be critical obstacles, like the huge volume of propellant tanks meant to contain liquid hydrogen with a density of just 70 kg/m^3 when liquid at near boiling point creating too much drag for spaceplanes to accelerate, to minor things like liquid hydrogen being more expensive than kerosene.

Many missions won’t even make use of hydrolox engines’ superior exhaust velocity. Low deltaV requirements means that performance is wasted. Long duration missions are also less suited to hydrolox propulsion, as storing liquid hydrogen against the heat of unfiltered sunlight beating down on gigantic propellant tanks is a real pain.

==Design and Function==
Hydrogen-oxygen engines are, like most chemical engines, fairly simple. You get some hydrogen, you get some oxygen, you pump them into a chamber, you introduce a spark or other ignition source, and steam comes out the back glowing with a very pretty clear flame, sometimes with varying minor coloration.

Hydrolox engines are almost universally fuel-rich and make great use of the excess hydrogen by using it as a coolant for regenerative cooling. This is when propellant flow is used to extract heat from the hottest parts of the engine, such as the nozzle throat, and then adding it into the reaction chamber to improve overall efficiency.

The biggest source of design variations is in how the propellant pumps are powered.

The J-2 used a gas generator, having a secondary burner to power the turbopumps. The J-2S cut that, by tapping off a portion of the main chamber gas flow, and avoiding a separate combustion chamber. You can get more complex, like the multi-stage booster pumps the SSME used to get to nearly 207 bars of pressure. Alternately, you can go dead simple--the expander cycle cleverly manages to harvest all the pumping energy it needs for low-pressure vacuum engine operation from the hydrogen used to cool the chamber and nozzle bell, meaning the turbine inlet gas is only mildly warm. That’s perfect if you want to use turbopump blades made out of a lightweight and strong material that can’t handle high temperature gases, like aluminium alloys.

==Performance / Capabilities / Applications==
The main benefit of hydrolox is its performance: even a moderate pressure engine with a vacuum nozzle can break 440 seconds of specific impulse. No other engine type in present-day service can reach those levels of performance. At the same time, hydrogen is a fantastic coolant, more than an order of magnitude better than any other fuel, meaning it’s possible to build some incredibly clever engine cycles without problems like turbine temperature or propellant “coking” of hydrocarbons.

As an example of the cleverness this allows, the turbine inlet temperature on the expander cycle RL-10 is below the boiling point of water, more like a kitchen appliance than the complex high-temperature turbomachinery needed in other fuels.

Similar endurance and robustness was seen with the J-2 engines for Saturn. The [i]Space Shuttle Decision[/i] by T A Heppenheirner quotes Rocketdyne's Paul Castenholz, one of the engineers managing J-2 development. "We never wore out an engine of the J-2 type. We could run it repeatedly; there was no erosion of the chamber, no damage to the turbine blades. If you looked at a J-2 after a hot firing, you would not see any difference from before that firing. The injectors always looked new; there was no erosion or corrosion on the injectors.” As Heppenheimer reports, this was inspired by examples like a single test engine which ran for 103 starts and 6.5 hours without overhaul. Higher performance hydrolox engines like the Space Shuttle Main Engine (SSME) have a more complex reputation, though more recent tests of the AR-22 variant demonstrated 10 firings in 10 days with no major overhaul. With good performance from even relatively simple engines, it’s no wonder that rocket pioneers like Werner Von Braun were swayed to the Cult of Hydrolox from initial skepticism. For more than half a century, it’s been the propellant of choice for upper and sustainer stages for most American launch vehicles from Atlas to Delta to the Space Shuttle. For that reason, if you want performance with near-term engines, hydrolox has a lot to recommend it.

The downside of hydrolox is the blasted hydrogen. Its performance comes from hydrogen’s low molecular mass--about the smallest anything can get and be legitimately called a molecule--but that also means it has to be incredibly cold to be liquid, at about 20 C colder than liquid oxygen. That’s troublesome for long duration storage, making “boiloff” the word of the day. Heat will get to your hydrogen any way it can, and before long, you won’t have any left. Even once kept liquid, the density is incredibly poor. That means heavier and physically larger tanks for the same performance. Even with hydrogen’s improved specific impulse, the Hydrolox Delta IV lower stage had to be five meters in diameter to lift the same payloads as the 3.8m Atlas V, since hydrolox has about a third the mixed density of kerolox.

==Worked Example(s)==
Planned to go here:

*RL-10 expander
*J-2S gas-gen
*SSME
*Maybe M-1 massive gas gen?

==Additional References==
GRAB SHUTTLE DECISION LINKS