Volume 4 fuel cells and hydrogen technology 4 02 – current perspective on hydrogen and fuel cells

Volume 4 fuel cells and hydrogen technology 4 02 – current perspective on hydrogen and fuel cells Volume 4 fuel cells and hydrogen technology 4 02 – current perspective on hydrogen and fuel cells Volume 4 fuel cells and hydrogen technology 4 02 – current perspective on hydrogen and fuel cells Volume 4 fuel cells and hydrogen technology 4 02 – current perspective on hydrogen and fuel cells

4.02.1.4.1 Wide-Field Infrared Survey Explorer

4.02.3 Other Current Uses of Hydrogen and Fuel Cells

4.02.3.1 Current Uses of Hydrogen

4.02.3.2 Current Uses of Fuel Cells

4.02.3.2.1 Electric power generation

4.02.3.2.2 Backup power supplies

4.02.1 Space Applications of Hydrogen

The general public in recent years has been exposed more and more to the topics of hydrogen and fuel cells, and with good reason Worldwide concern about the diminishing supply of hydrocarbon fuels that play a fundamental role in the world energy mix and an equally worldwide concern over the effects on the world’s environment of burning hydrocarbon fuels are two compelling reasons why hydrogen and fuel cells are frequent topics in the world’s daily news

Even though hydrogen is produced in great quantities industrially, the general public has little or no direct experience with hydrogen,

so the public’s general knowledge is limited to film footage of the Hindenburg burning and crashing and rockets violently leaving the launch pad The gap in knowledge is filled with instinctual suspicion and fear of the unknown This current public perspective on hydrogen is changing as more information about hydrogen and a ‘hydrogen economy’ is available Over time, the image of hydrogen as

a future potential fuel in the public mindset is becoming less fearsome and more appreciated for its unique and beneficial characteristics Hydrogen is the ‘cleanest’ of all fuels because as it is oxidized (burned), it produces only water Hydrogen is abundant over the entire planet, but because of hydrogen’s reactivity, it is rarely found in its pure gaseous state Hydrogen is readily produced from methane, or by the electrolysis of water For its mass, hydrogen packs a lot of energy; a fact that makes it a highly used fuel for rockets The public is largely unaware of how critical hydrogen is to the production of commonly used commodities such as gasoline and fertilizer This is probably because the hydrogen is produced where it is used, at industrial sites well beyond public view and awareness Hydrogen today is largely produced from natural gas which is primarily obtained from nonrenewable sources as are other hydrocarbon fuels

One of hydrogen’s uses that has been widely appreciated is its use for space travel Hydrogen has been critical to both the propulsion of spacecraft and the generation of electrical power while in space Hydrogen is widely used by the United States,

Oxidizer Fuel fuel to oxidizer (s, sea level) (kg-s l−1, SL)

4.02.1.1 Space Propulsion

Hydrogen use as a chemical propellant is unquestionably the most important space application of hydrogen In terms of mass of hydrogen used, the use of hydrogen propellant dwarfs all other space applications of hydrogen combined The reason for the use of hydrogen is that hydrogen is the most efficient propellant Specific impulse is a measure of the change of momentum per unit weight of the propellant on Earth The higher the specific impulse, the less the weight of propellant needed Table 1 shows a comparison of liquid fuel rocket propellants [1]

4.02.1.1.1 Atlas–Centaur

The Atlas–Centaur rocket was the first rocket to use the combination of liquid hydrogen/oxygen (LH2/LOX) for propulsion This rocket’s second stage, Centaur, used the RL10 LH2/LOX rocket engine shown in Figure 1 The RL10 manufactured by Pratt and Whitney was the first engine to use LH2/LOX Figure 2 shows a simplified flow schematic of the RL10 engine The RL10 used liquid hydrogen to cool the engine nozzle, and the heat absorbed by the liquid hydrogen caused the hydrogen to expand, after which it flowed through a turbine The rotation of the turbine was mechanically coupled to the LH2 and LOX pumps which pump the propellants to the combustion chamber This synergy of design made the engine lightweight and very reliable The RL10 was first run in July 1959, and was first flight tested with the Centaur on 27 November 1963 The test flight was delayed to allow for the mourning of President Kennedy, who was shot days earlier [4] Upgraded versions of the RL10 are used to currently launch Atlas–Centaur rockets and on the upper stage of the Delta IV rocket being used currently

Figure 3 shows a cutaway view of the Centaur rocket The Centaur rocket uses helium to pressurize the hydrogen and oxygen propellants The pressurized propellants rigidize the structure of the rocket which minimizes the mass of the rocket structure Using pressurized propellants to rigidize a rocket’s structure was initially demonstrated during a US Air Force missile project called MX-774 Although this project was cancelled before the development was completed, Charlie Bossart, the MX-774 designer, applied this knowledge to the design of the Atlas Later, this ‘pressure-stabilized’ approach was used by Krafft Ehricke as a key design element

of the Centaur [4] Figure 4 shows the dual engine version of the Centaur upper stage

Together, the RL10 LH2/LOX rocket engine and the lightweight Centaur structure gave NASA a lightweight powerful upper stage The powerful Centaur upper stage when mated to the lightweight Atlas provided NASA with a launch vehicle capable of putting payloads beyond low Earth orbit (LEO; 160–2000 km altitude) and out into space (>2000 km altitude) Figure 5 shows an exploded view of an Atlas–Centaur rocket

4.02.1.1.2 Apollo Saturn

The Apollo missions to the moon used the Saturn V rocket The Saturn V was a three-stage rocket The first stage was to carry the Saturn rocket to an altitude of ∼200 000 feet (61 km) It used five engines that burned kerosene and liquid oxygen as propellant to produce 7 600 000 lbs (34 000 000 N) of thrust There were no LH2/LOX engines then (or now) capable of producing such enormous thrust Its second and third stages which burned sequentially used LH2/LOX The payloads to be launched required more powerful LH2/LOX engines (200 000 lbs/890 000 N of thrust) than the RL10 LH2/LOX engine (20 000 lbs/89 000 N of thrust) This required the development of the J-2 LH2/LOX by Rocketdyne shown in Figure 6

Figure 1 Pratt and Whitney RL10 [2]

LH2 pump Turbine LOX pump

Control valves

Heat exchanger

Combustion Chamber

Nozzle

Figure 2 RL10A-3 rocket engine flow schematic [3]

The second stage of the Saturn V, called SII, used five J-2 engines to provide 1 million pounds of thrust The SII stage had a diameter of 10 m and a length of ∼24.9 m Filled with propellants, its gross mass was ∼480 000 kg, and when empty, a mass of

∼36 000 kg The burn time for the second stage was 367 s [9] Figure 7 shows an illustration of the SII stage, and Figure 8 shows the SII stage being hoisted into a test stand at the NASA Stennis Space Center

The third stage of the Saturn V used one J-2 engine, which provided ∼200 000 lbs of thrust The SIII stage had a diameter of 6.6 m and a length of ∼17.8 m Filled with propellants, its gross mass was 119 900 kg, and when empty, a mass of ∼11 000 kg The burn time for the second stage was 475 s [12]

An upgraded version of the J-2 engine, called J-2X, is being currently developed for NASA for further exploration of the Moon and Mars The J-2X with 294 000 lbs of thrust will be more powerful than the J-2 engine The J-2X engine requirements call for an exit

Stub adapter LH2 tank Fuel slosh

shield

Intermediate bulkhead Propellant utilization probeLOX tank

Intermediate adapter

The Space Shuttle was launched using a combination of two solid rocket motors and three LH2/LOX engines The solid rocket motors burn a solid propellant and cannot adjust their level of thrust, whereas the LH2/LOX engines use liquid propellants and

Figure 3 Cutaway view of Centaur upper stage [5]

Centaur interstage adapter Booster interstage adapter

RD-180 engine

Vehicle effectivity

THRU SA-203

SA-204 THRU SA-207&SA-501 THRU SA-503

SA-208 &

subsequent;

and SA-504 &

subsequent Thrust (altitude) 200 000 lb 225 000 lb 230 000 lb

Specific impulse

Figure 5 Atlas–Centaur rocket [7]

Figure 6 Rocketdyne J-2 rocket engine From NASA Image Exchange, ID No MSFC-9801770, NASA Marshall Space Flight Center (J2 Engine Slide) http://mix.msfc.nasa.gov/IMAGES/MEDIUM/9801770.jpg [8]

have their thrust adjusted during the ascent Unlike the Saturn V where the first stage is nearly complete with its burn when the second stage starts its burn, the Space Shuttle’s propulsion combination propelled the Space Shuttle simultaneously until the solid rocket motors completed their burn The solid rocket motors then separated from the Space Shuttle and fell back to Earth where they were recovered, refurbished, and reused The three LH2/LOX engines then continued to burn, eventually lifting the Space Shuttle into LEO Figure 9 shows the Space Shuttle shortly after liftoff with both the solid rocket motors and the LH2/LOX engines operating The large orange tank shown in Figure 9 is the LH2/LOX storage tank This tank was the only expendable portion of the Space Transportation System, separating from the Space Shuttle after the LH2/LOX engines completed their burn Unlike the Saturn

V and all previous rockets, the Space Shuttle was reusable Following its 1–2-week mission, the Space Shuttle reentered the Earth’s atmosphere, glided, and then landed on a landing strip

LH2 Tank Manhole cover pressure line

Cable tunnel Gas distributor

LOX vent line Mast Second

Fuel level sensor J-2 engine

Heat shield Ring slosh baffle

LOX sump

LH2 suction line

Ullage rocket Saturn V

Figure 7 Saturn V second stage, SII From Marshall Space Flight Center Image Exchange, Photo No 9801810 http://mix.msfc.nasa.gov/IMAGES/ MEDIUM/9801810.jpg [10]

Figure 8 SII lifted into test stand From NASA Image Exchange, ID No 67-701-c, NASA Stennis Space Center (SII stage of Saturn V rocket photo) http:// www.ssc.nasa.gov/sirs/photos/history/low/67-701-c.jpg [11]

The Space Shuttle Main Engine (SSME) was unique, in that it was the only reusable LH2/LOX engine in the world It was designed for 7.5 h of operation over an average life span of 55 starts [15] The SSME was much more powerful than the J-2 engine (470 000 lbs of thrust (at vacuum) vs 200 000 lbs) The SSME consumed 4420 kg per minute of LH2 and 26 450 kg per minute of LOX for ∼8.5 min The SSME had an exit diameter of 230 cm, a length of 427 cm, and a mass of 3526 kg [16, 17] Figures 10 and 11 show the SSME [18, 19] The SSME is one of the rocket engines being considered for the first stage of NASA’s new SLS

4.02.1.1.4 Delta IV

The Delta family of rockets started with the first launch of a Delta rocket on 13 May 1960 The early Delta rockets did not use LH2 as

a propellant The Delta rocket evolved through continuous improvement from a Delta A rocket in 1962 through a Delta N rocket in

Figure 9 Space Shuttle after Liftoff From NASA Image Exchange, ID No STS062(S)055, NASA Johnson Space Center http://images.jsc.nasa.gov/lores/ STS062(S)055.jpg [14]

Figure 10 Space Shuttle Main Engine From NASA Image Exchange, ID No KSC-04PD-1643, NASA Kennedy Space Center http://images.ksc.nasa.gov/ photos/2004/low/KSC-04PD-1643.gif [18]

Figure 11 Test firing of the SSME From NASA Image Exchange, Photo No MSFC-7995081 http://mix.msfc.nasa.gov/IMAGES/

The RS-68 LH2/LOX rocket engine was developed by Rocketdyne in 1998 The RS-68 is the largest and most powerful LH2/LOX engine in the world The RS-68 is more powerful than the SSME engine (758 000 lbs of thrust vs 470 000 lbs) The RS-68 is ∼244 cm

in diameter and 521 cm in length, and has a mass of 6761 kg [23] Figure 12 shows the RS-68 being test fired [24].The RS-68 is a candidate for possible use in NASA’s new SLS

4.02.1.1.5 Other LH2/LOX-powered rockets

Only recently, nations other than the United States have started to produce LH2/LOX engines The Japanese have produced two engines which are used for first and second stages The LE-5 is used for the second stage, and had its first flight in 1986 The LE-5 produces 23 100 lbs of thrust and has a diameter of 2.49 m, a length of 2.68 m, and a mass of 245 kg [25] Figure 13 shows the LE-5 [26] Japan also produced the LE-7 designed for a first stage, and had its first flight in 1994 The LE-7 produces 242 000 lbs of thrust and has a diameter of 4 m, a length of 3.4 m, and a mass of 1714 kg [27] Figure 13 also shows the LE-7 [26] Japan’s H-II unmanned expendable launch vehicle uses one LE-7A engine in its first stage and one LE-5 is used on the upper stage

Snecma Moteurs (France) has produced the Vulcain, an LH2/LOX engine designed for a first stage, and had its first successful flight in 1997 [28] The Vulcain produces 256 760 lbs of thrust and has a diameter of 1.76 m, a length of 3.0 m, and a mass of

1700 kg [29] Figure 14 shows the Vulcain [28] The Ariane 5 launch vehicle uses a single Vulcain as its first stage engine The Energia–Buran shown in Figure 15(a) is a Russian-built launch system whose second stage (shown in Figure 15(b)) is fueled

by LH2/LOX The second stage is powered by four RD-0120 engines shown in Figure 16 The RD-0120 was similar in performance

to the US-built SSME used in the US Space Shuttle The Energia flew only two flights The first of these was on 15 May 1987 with Polyus spacecraft as the payload The second and final launch was with the Russian shuttle vehicle, Buran, on 15 November 1988 The fall of the Soviet Union ended the flights of the Energia and the use of the RD-0120 engines In the 1990s, several other launchers were designed with an Energia core stage, but none were ever built Starting in 2001, another Russian LH2/LOX engine, the RD-0146, was developed, but as of 2010 had not yet flown

Figure 12 RS-68 test firing From NASA Image Exchange, Photo No MSFC-0700063 http://www.nasa.gov/images/content/

148709main_d4_testing_08.jpg [24]

LE-5/5A/5B (LOX/LH)

LE-7 (LOX/LH)

Figure 13 Japan’s LE-7 LH2/LOX engine [25]

4.02.1.1.6 Advanced space propulsion technology

Unlike all previous mentioned rocket engines that use hydrogen, advanced space propulsion that uses hydrogen does not combust the hydrogen Instead, the hydrogen is heated to high temperatures from high-energy sources and then the hydrogen exits the engine

at a high velocity The specific impulse from these engines varies depending on the temperature to which the hydrogen is heated, but

in general the specific impulse is much higher than that of conventional hydrogen/oxygen engines One common drawback to this

Figure 14 Vulcaine rocket engine [28]

Figure 15 (a) Energia–Buran (b) Energia second stage [30, 31]

Figure 16 RD-0120 rocket engine [32]

propulsion technology is that a high-power, high-energy source is needed to heat the hydrogen There are several types of advanced rocket engines that take this approach

Hydrogen arcjet engines use an electric arc to directly heat the hydrogen which is then expanded through a rocket nozzle The engines have been operated with electric power sources ranging from 0.5 to 30 kW and have a characteristic specific impulse in the range of 1000–1500 s [33] Unlike chemical rockets which burn powerfully for short periods of time, these engines operate by providing small amounts of thrust over long periods of operation These engines minimize the amount of propellant needed for an overall mission, but are unsuitable for lifting rockets from the Earth’s surface Instead, these engines provide efficient propulsion after the payload is already in space This technology was investigated in the 1960s and more recently in the 1990s, and has been considered for orbital transfer and for satellite attitude control and orbit maintenance

Nuclear thermal rocket (NTR) engines use heat from a nuclear reactor to heat the hydrogen which is then expanded through a rocket nozzle In 1959, the first ground test of NTR technology was the test of ‘Kiwi-A’, a proof-of-concept test engine named after the New Zealand flightless bird [34, 35] Nuclear Engine for Rocket Vehicle Application (NERVA), shown in Figure 17, was developed during the 1960s as an upper stage engine to the Apollo Saturn V booster The initial test run of the engine was in September 1964 The engine had a specific impulse of 850 s and produced a thrust of 75 000 lbf A more powerful derivative, Phoebus, was capable of producing a thrust of 250 000 lbf In 1972, the development of a flight test model of the NERVA was cancelled when US plans for manned Mars exploration were cancelled by President Nixon

The Variable Specific Impulse Magnetoplasma Rocket (VASIMR) ionizes and heats a gas into a hot plasma using radiofrequency energy The heated plasma is then focused and directed using magnetic fields to generate thrust as shown in Figure 18 VASIMR is capable of using different gases (including hydrogen) VASIMR has demonstrated a very high-specific impulse of 4000 to over

10 000 s depending on propellant used [38], indicating that 90% or more of the propellant weight could be saved, a tremendous advantage, especially for long space voyages Like the hydrogen arcjet, it provides a low amount of thrust over a long period of time One serious drawback to this promising technology is that the VASIMR requires a very high-power electric source, and the mass of the power source offsets its propellant advantage Long planetary trips to date have used electric power sources that are much less powerful than what is needed by VASIMR Future, possibly very high-power nuclear electric sources, may allow this technology to be fully exploited Figure 19 shows an artist’s illustration of a VASIMR-powered spacecraft concept

4.02.1.2 Space Battery Power and Energy Storage – NiH2 Batteries

Nickel–hydrogen batteries were developed to increase energy density and capacity in rechargeable battery technology for aerospace energy storage The nickel–hydrogen cells are a hybrid technology, combining elements from both batteries and fuel cells The nickel–hydrogen cells utilize the nickel hydroxide electrode from nickel–cadmium cells and a platinum hydrogen electrode from

Shield Pressure vessel Reflector

NERVA reactor based on NRX A1 Radial

support Reactor

Graphite felt lateral support Top-loaded core

50 kW VASIMR laboratory experiment

Magnetic coils confine the ionized plasma

Helicon antenna ionizes gas to plasma

Quartz tube confines gas

Gaseous H2 or He Injection

Figure 17 NERVA rocket engine From NASA Image Exchange, Photo No MSFC–9902054, http://mix.msfc.nasa.gov/abstracts.php?p=1812 [36]

Figure 18 VASIMR engine test [37]

fuel cell technology to create a chemistry without the issues and limitations inherent with the cadmium electrode The nickel– hydrogen chemistry has better long-term cycle life and specific energy over the standard aerospace nickel–cadmium battery, though having a poorer volumetric energy density

The electrochemical reactions that occur during discharge are as follows:

NiOOH þ H2O þ e− → Ni OHÞð 2 þ OH− ðat the nickel electrodeÞ

1 =2 H2 þ OH− → H2O þ e− ðat the hydrogen electrodeÞ

1NiOOH þ =2 H2 → Ni OHÞð 2 ðoverall reaction during dischargeÞ

Figure 19 VASIMR-powered rocket concept [39]

a nickel–hydrogen cell is 1.25 VDC, about the same as for a nickel–cadmium cell

The development of nickel–hydrogen cells was started by COMSAT Laboratories in 1970 [40] After the initial demonstration of the feasibility of the nickel–hydrogen cell, INTELSAT funded COMSAT Laboratories to develop a 50 A-h cell, and in 1975, this development had progressed to the point that the US Naval Research Laboratory funded COMSAT Laboratories to develop a 35 A-h nickel–hydrogen cell for use on the US Navy’s Navigation Technology Satellite (NTS-2) spacecraft shown in Figure 20 [40, 42] The NTS-2, launched in

1977, was the first use of nickel–hydrogen battery technology in space Nickel–hydrogen cells were then put in service on Intelsat V, VI, and VII satellites from 1983 through 1996 [43, 44] Separate designs evolved to address LEO operations In 1990, the Hubble telescope launched as the first LEO satellite using nickel–hydrogen batteries, after which nickel–hydrogen batteries were used for numerous LEO missions In the mid 1980s, the International Space Station (ISS) power system was designed with the largest ever series-connected nickel–hydrogen battery Orbital Replacement Units (ORUs), as shown in Figure 21, to provide energy storage during the LEO eclipse period [45] The first set of ISS battery ORUs was launched in 2000 The ISS nickel–hydrogen batteries have an operational life of 6.5 years

Graphitecell sleeve

Fuse assembly

Battery signalconditioning andcontrol module Status indicator

Acme screw

Multiple layer insulation EVA tether

Nickel hydrogen

battery cell

Figure 21 ISS NiH2 battery ORU [45]

Early nickel–hydrogen cells were packaged in individual pressure vessels (IPVs) shown in Figure 20 [41].Subsequent develop­ment packaged multiple cells within a single pressure vessel (SPV) These designs addressed the low-energy density of nickel–hydrogen systems These were known as common pressure vessel (CPV) or SPV designs

Nickel–hydrogen cell technology started to become obsolete in the 2000s with the introduction of lithium-ion battery technology that had significantly better specific energy, energy density, lower self-discharge rates, and higher columbic efficiency Unlike nickel–hydrogen, which was almost exclusively an aerospace technology, the commercial development of lithium-ion technology for small portable electronics provided the added push, advancing lithium-ion technology and accelerating the displacement of nickel–hydrogen technology in space

4.02.1.3 Astronaut Environmental Control and Life Support

As manned space missions have gradually increased in number of crew and length in time, the advantage of recycling life supporting resources such as water and oxygen versus supplying these resources as expendable (not recycled) becomes more and more compelling when comparing the mass of spacecraft using recycling system with the alternative expendable systems The disadvan­tage to recycling systems is that they generally are not as reliable as expendable systems; they also tend to consume another valuable resource, electrical power Water electrolysis equipment that contributes to ‘closing the oxygen loop’ has been developed and is being used on the ISS Water electrolysis splits water into hydrogen and oxygen The oxygen contained within the water is recycled, while the hydrogen is either vented overboard into space or used to reduce carbon dioxide The reduction of carbon dioxide with hydrogen produces water and either methane or amorphous carbon black The carbon dioxide reduction process that reduces carbon dioxide to water vapor and methane is called the Sabatier CO2 reduction system [46] Another alternative carbon dioxide reduction process that reduces carbon dioxide to water vapor and carbon black is called the Bosch CO2 reduction system The water vapor produced during the reduction process is subsequently recovered and split into hydrogen and oxygen by water electrolysis A block diagram of the advanced life support system is shown in Figure 22

4.02.1.4 Scientific Instrument Cooling

Cooling technologies are required for the high-performance detection of electromagnetic radiation of millimeter to nanometer wavelengths Radiation detectors are subject to background noise that unless substantially reduced can make precise detection impossible Radiation detectors are cooled to cryogenic temperatures to reduce the background noise, sometimes referred to as the background limit, caused by thermal energy of the detectors themselves This applies for many types of radiation detectors For infrared detectors looking through the Earth’s atmosphere, the background noise is frequently determined by the atmosphere, but

H2O

heat exchanger

Process atmosphere

Potable water processor

Potable water storage

Figure 22 Advanced environmental control and life support for space [47]

for infrared instruments in space, above the atmosphere, the background limit is determined by the instrument itself which glows with thermal infrared radiation These infrared instruments are cooled to less than 1 K To provide such low-temperature cooling requires a refrigerant capable of working at these low temperatures, typically helium or hydrogen

4.02.1.4.1 Wide-Field Infrared Survey Explorer

On 14 December 2009, NASA launched the Wide-Field Infrared Survey Explorer (WISE) The WISE, shown in Figure 23, circles the Earth, from pole to pole, 15 times each day, taking pictures of the sky every 11 s Its mission is to observe the universe in the 3–25 μm wavelength [49] The WISE cryostat, shown in Figure 24, is a two-stage, solid hydrogen cryostat that cools the satellite’s detectors to 7.6 K The solid hydrogen cryostat is expected to last about 10 months after which the satellite’s mission will be finished By cooling the optics and detectors to such a low temperature, WISE will be able to measure the infrared glow of celestial objects with a sensitivity hundreds of times more than any previous radiometer WISE will be able to measure the infrared radiation of asteroids in our solar system between Mars and Jupiter and provide the first good estimate of their size distribution

Figure 24 WISE hydrogen cryostat From Jet Propulsion Laboratory Photo Journal, Photo No PIA12316 http://photojournal.jpl.nasa.gov/jpegMod/ PIA12316_modest.jpg [50]

Figure 25 Planck spacecraft [51]

Bang creation of the universe These minute differences are measured in millionths of a degree Kelvin about the observable temperature difference measured from Earth of the heat of a rabbit sitting on the moon [52] To obtain this extraordinary sensitivity, the radiation detectors must be cooled to 0.1°K The cooling system for the Planck spacecraft uses a three-stage cooler The second cooling stage of the Planck spacecraft shown in Figure 26 uses a hydrogen sorption cooler Gaseous hydrogen at low pressure is absorbed by a metal hydride sorbent bed When the metal hydride sorbent bed has absorbed the hydrogen, the bed is then heated to release the hydrogen, creating high pressure hydrogen The high pressure hydrogen is precooled, then expanded through a Joule–Thompson expander which cools the detectors to below 20 K The expanded hydrogen is reabsorbed by the metal hydride sorbent bed so that it can be recompressed at the start of another cooling cycle The heat switch is activated to cool the sorbent bed after the bed has been heated to release the hydrogen Cooling the sorbent bed prepares it to reabsorb hydrogen at the start of the next cooling cycle

Gaseous hydrogen at low P

Liquid refrigerant J-T

Heat from sensors

Figure 26 Planck cooling system [53]

4.02.2 Space Applications of Fuel Cells

Fuel cell technology was put into the world public spotlight in dramatic fashion by NASA which chose to use hydrogen/oxygen fuel cells to provide critical electrical power to its manned spacecraft While earlier development efforts were looking at fuel cells for various terrestrial power applications, these efforts gained little public notice, but the world public watched as astronauts were launched into space, voyaged to the moon, and returned to Earth, powered in no small measure by a mysterious, new ‘space-age’ technology called fuel cells NASA first used fuel cells for powering the Gemini spacecraft during takeoff and while in orbit Besides providing the electricity for the Gemini spacecraft, fuel cells were also used to provide power for the Apollo spacecraft and the Space Shuttle The choice of fuel cells as the power source was because fuel cells offered the lightest electrical power source Also the water produced by the fuel cells used for the Apollo and the Space Shuttle was suitable for astronauts to use which saved the weight of bringing a separate source of water For each spacecraft, the reactant use was ∼0.33 kg O2h−1 and 0.04 kg H2 h−1 per kilowatt of power produced The water generation was ∼0.37 kg h−1 per kilowatt of power produced

4.02.2.1 Gemini

On 21 August 1965, the Gemini 5 spacecraft launched the use of hydrogen–oxygen fuel cells for electrical power generation The use

of fuel cells was still in doubt as late as November 1963 when fuel cell production was stopped due to technical problems In January 1964, a meeting at the Johnson Space Center was held to review the development status and decide what to do It was decided to redesign the fuel cells and have them ready for the fifth Gemini flight [54] The Gemini fuel cell system was used on Gemini 5, 7, 8, 9A, 10, 11, and 12 The Gemini fuel cells were constructed with a polymer membrane which served as the ion-conductive electrolyte between the cells anode and the cathode plates This technology served as the precursor for the proton exchange membrane (PEM) fuel cells that are being developed for many of the fuel cell terrestrial applications that are currently under development

The Gemini fuel cell stack, shown in Figure 27, consisted of 32 cells in electrical series Each cell contained an active area of

360 cm2 (20 cm 18 cm) and used an acid-based PEM as its electrolyte Three fuel cell stacks were connected electrically in parallel, shown in Figure 28, and housed in a fuel cell section shown in Figure 29 Each section was about 66 cm in length and 33 cm in diameter and weighed about 31 kg and produced a maximum of 600 W Two such fuel cell sections plus an associated reactant supply system comprise the Gemini fuel cell system shown in Figure 30

Hydrogen was distributed to each section, and within each section, the hydrogen was manifolded to each of the three cell stacks Oxygen was distributed to each section and within each section it was manifolded to each cell stack The oxygen compartments of the cells were open to the inside of the section container which was filled with oxygen The hydrogen and oxygen were not circulated through the cell stacks, but were ‘dead-ended’ and kept at a constant pressure so that as hydrogen and oxygen was consumed by the cells, more hydrogen and oxygen was supplied to the cells to maintain the constant pressure The water produced by each cell was absorbed by a wick immediately adjacent to the oxygen compartment of each