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==History== | ==History== | ||
[[File:RIAN archive 303890 A battery of Katyusha during the 1941-1945 Great Patriotic War.jpg|thumb|A battery of [[Katyusha rocket launcher]]s fires at German forces during the [[Battle of Stalingrad]], 6 October 1942]] | |||
The medieval [[Song dynasty]] Chinese invented a very primitive form of solid-propellant rocket.<ref>{{Cite book |title=Space Science in China |last= Hu |first=Wen-Rui |year=1997 |isbn= 978-9056990237 |publication-date=August 20, 1997 |page=15}}</ref> Illustrations and descriptions in the 14th century Chinese military treatise ''[[Huolongjing]]'' by the Ming dynasty military writer and philosopher [[Jiao Yu]] confirm that the Chinese in 1232 used proto solid propellant rockets then known as "[[fire arrows]]" to drive back the Mongols during the [[Mongol siege of Kaifeng]].<ref name="Greatrix 2012 1">{{Cite book |title=Powered Flight: The Engineering of Aerospace Propulsion |last=Greatrix |first= David R. |publisher=Springer |year=2012 |isbn=978-1447124849 |pages=1}}</ref><ref name="Nielsen 1997 2–4">{{Cite book |title=Blast Off!: Rocketry for Elementary and Middle School Students P |last= Nielsen |first= Leona |publisher= Libraries Unlimited |year=1997 |isbn= 978-1563084386 |pages=2–4}}</ref> Each arrow took a primitive form of a simple, solid-propellant rocket tube that was filled with gunpowder. One open end allowed the gas to escape and was attached to a long stick that acted as a guidance system for flight direction control.<ref name="Nielsen 1997 2–4"/><ref name="Greatrix 2012 1"/> | The medieval [[Song dynasty]] Chinese invented a very primitive form of solid-propellant rocket.<ref>{{Cite book |title=Space Science in China |last= Hu |first=Wen-Rui |year=1997 |isbn= 978-9056990237 |publication-date=August 20, 1997 |page=15}}</ref> Illustrations and descriptions in the 14th century Chinese military treatise ''[[Huolongjing]]'' by the Ming dynasty military writer and philosopher [[Jiao Yu]] confirm that the Chinese in 1232 used proto solid propellant rockets then known as "[[fire arrows]]" to drive back the Mongols during the [[Mongol siege of Kaifeng]].<ref name="Greatrix 2012 1">{{Cite book |title=Powered Flight: The Engineering of Aerospace Propulsion |last=Greatrix |first= David R. |publisher=Springer |year=2012 |isbn=978-1447124849 |pages=1}}</ref><ref name="Nielsen 1997 2–4">{{Cite book |title=Blast Off!: Rocketry for Elementary and Middle School Students P |last= Nielsen |first= Leona |publisher= Libraries Unlimited |year=1997 |isbn= 978-1563084386 |pages=2–4}}</ref> Each arrow took a primitive form of a simple, solid-propellant rocket tube that was filled with gunpowder. One open end allowed the gas to escape and was attached to a long stick that acted as a guidance system for flight direction control.<ref name="Nielsen 1997 2–4"/><ref name="Greatrix 2012 1"/> | ||
The first rockets with tubes of cast iron were used by the [[Kingdom of Mysore]] under [[Hyder Ali]] and [[Tipu Sultan]] in the 1750s. These rockets had a reach of targets up to a mile and a half away. These were extremely effective in the [[Second Anglo-Mysore War]] that ended in a humiliating defeat for the [[British Empire]]. Word of the success of the Mysore rockets against the British Imperial power triggered research in England, France, Ireland and elsewhere. When the British finally conquered the fort of [[Siege of Seringapatam (1799)|Srirangapatana]] in 1799, hundreds of rockets were shipped off to the [[Royal Arsenal]] near London to be reverse-engineered. This led to the first industrial manufacture of military rockets with the [[Congreve rocket]] in 1804.<ref name="Bowdoin 2004">{{Cite book |title=Rockets and Missiles: The Life Story of a Technology|last= Van Riper |first= Bowdoin |publisher= The Johns Hopkins University Press |year=2004 |isbn= 978-0801887925 |pages=14–15}}</ref> | The first rockets with tubes of cast iron were used by the [[Kingdom of Mysore]] under [[Hyder Ali]] and [[Tipu Sultan]] in the 1750s. These rockets had a reach of targets up to a mile and a half away. These were extremely effective in the [[Second Anglo-Mysore War]] that ended in a humiliating defeat for the [[British Empire]]. Word of the success of the Mysore rockets against the British Imperial power triggered research in England, France, Ireland and elsewhere. When the British finally conquered the fort of [[Siege of Seringapatam (1799)|Srirangapatana]] in 1799, hundreds of rockets were shipped off to the [[Royal Arsenal]] near London to be reverse-engineered. This led to the first industrial manufacture of military rockets with the [[Congreve rocket]] in 1804.<ref name="Bowdoin 2004">{{Cite book |title=Rockets and Missiles: The Life Story of a Technology|last= Van Riper |first= Bowdoin |publisher= The Johns Hopkins University Press |year=2004 |isbn= 978-0801887925 |pages=14–15}}</ref> | ||
In 1921 the [[Soviet]] research and development laboratory [[Gas Dynamics Laboratory]] began developing solid-propellant rockets, which resulted in the first launch in 1928, which flew for approximately 1,300 metres.<ref name="RSB_GDL">{{cite web |last1=Zak |first1=Anatoly |title=Gas Dynamics Laboratory |url=http://www.russianspaceweb.com/gdl.html |website=Russian Space Web |access-date=29 May 2022}}</ref> These rockets were used in 1931 for the world's first successful use of rockets to [[JATO|assist take-off of aircraft]] were carried out<ref name="Glushko">{{cite book |last1=Glushko |first1=Valentin |title=Developments of Rocketry and Space Technology in the USSR |date=1 January 1973 |publisher=Novosti Press Pub. House |page=7 |url=https://www.amazon.com/Development-rocketry-space-technology-USSR/dp/B0006CHI4I}}</ref> and became the prototypes for the [[Katyusha rocket launcher]],<ref name="Ezo_Katyusha">{{cite web |title=Katyusha rocket launcher. Weapon of Victory: multiple launch rocket system "Katyusha" |url=https://ezoteriker.ru/en/reaktivnyi-minomet-katyusha-oruzhie-pobedy-reaktivnaya-sistema-zalpovogo-ognya/ |website=ezoteriker |access-date=5 June 2022}}</ref> which were used during [[World War II]]. | |||
In the [[United States]] modern castable composite solid rocket motors were invented by the American aerospace engineer [[Jack Parsons (rocket engineer)|Jack Parsons]] at [[Caltech]] in 1942 when he replaced double base propellant with roofing [[asphalt]] and [[potassium perchlorate]]. This made possible slow-burning rocket motors of adequate size and with sufficient shelf-life for jet-assisted take off applications. [[Charles Bartley]], employed at JPL (Caltech), substituted curable [[synthetic rubber]] for the gooey asphalt, creating a flexible but geometrically stable load-bearing propellant grain that bonded securely to the motor casing. This made possible much larger solid rocket motors. Atlantic Research Corporation significantly boosted composite propellant I<sub>sp</sub> in 1954 by increasing the amount of powdered aluminium in the propellant to as much as 20%.<ref>{{cite book |author= M. D. Black |title= The Evolution of Rocket Technology |page= 39 |publisher= Native Planter, SLC |date= 2012 }} payloadz.com under ''ebook/History'' {{dead link|date= July 2014}}</ref> | |||
Solid-propellant rocket technology got its largest boost in technical innovation, size and capability with the various mid-20th century government initiatives to develop increasingly capable military missiles. After initial designs of [[ballistic missile]] military technology designed with [[liquid-propellant rocket]]s in the 1940s and 1950s, both the [[Soviet Union]] and the [[Federal government of the United States|United States]] embarked on major initiatives to develop solid-propellant [[Tactical ballistic missile|local]], [[Medium-range ballistic missile|regional]], and [[Intercontinental ballistic missile|intercontinental]] ballistic missiles, including solid-propellant missiles that could be launched from [[Air-launched ballistic missile|air]] or [[Submarine-launched ballistic missile|sea]]. Many [[List of missiles by country|other governments]] also developed these military technologies over the next 50 years. | Solid-propellant rocket technology got its largest boost in technical innovation, size and capability with the various mid-20th century government initiatives to develop increasingly capable military missiles. After initial designs of [[ballistic missile]] military technology designed with [[liquid-propellant rocket]]s in the 1940s and 1950s, both the [[Soviet Union]] and the [[Federal government of the United States|United States]] embarked on major initiatives to develop solid-propellant [[Tactical ballistic missile|local]], [[Medium-range ballistic missile|regional]], and [[Intercontinental ballistic missile|intercontinental]] ballistic missiles, including solid-propellant missiles that could be launched from [[Air-launched ballistic missile|air]] or [[Submarine-launched ballistic missile|sea]]. Many [[List of missiles by country|other governments]] also developed these military technologies over the next 50 years. | ||
By the later 1980s and continuing to 2020, these government-developed highly-capable solid rocket technologies have been applied to [[orbital spaceflight]] by many<!-- possibly all of the governments of countries that have developed orbital spaceflight technology, which is 6 or 8 now; but would need to confirm --> [[nation state|government-directed programs]], most often as [[solid rocket booster|booster rockets]] to add extra thrust during the early ascent of their primarily liquid rocket [[launch vehicle]]s. Some designs have had solid rocket upper stages as well. Examples flying in the 2010s include the | By the later 1980s and continuing to 2020, these government-developed highly-capable solid rocket technologies have been applied to [[orbital spaceflight]] by many<!-- possibly all of the governments of countries that have developed orbital spaceflight technology, which is 6 or 8 now; but would need to confirm --> [[nation state|government-directed programs]], most often as [[solid rocket booster|booster rockets]] to add extra thrust during the early ascent of their primarily liquid rocket [[launch vehicle]]s. Some designs have had solid rocket upper stages as well. Examples flying in the 2010s include the European [[Ariane 5]], US [[Atlas V]] and [[Space Shuttle Solid Rocket Booster|Space Shuttle]], and Japan's [[H-II]]. | ||
The largest solid rocket motors ever built were Aerojet's three {{convert|260|in|m|adj=on|sigfig=3|order=flip|sp=us}} monolithic solid motors cast in Florida.<ref>{{Cite journal|url = https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20000033816.pdf|title = The 260 - The Largest Solid Rocket Motor Ever Tested|date = June 1999|access-date = July 24, 2014|website = nasa.gov}}</ref> Motors 260 SL-1 and SL-2 were {{convert|261|in|m|sigfig=3|order=flip|sp=us}} in diameter, {{convert|80|ft|8|in|m|order=flip|sp=us}} long, weighed {{convert|1,858,300|lb|kg|order=flip}}, and had a maximum thrust of {{convert|3.5|e6lbf|MN lbf|order=out|abbr=on}}. Burn duration was two minutes. The nozzle throat was large enough to walk through standing up. The motor was capable of serving as a 1-to-1 replacement for the 8-engine [[Saturn I]] liquid-propellant first stage but was never used as such. Motor 260 SL-3 was of similar length and weight but had a maximum thrust of {{convert|5.4|e6lbf|MN lbf|order=out|abbr=on}} and a shorter duration. | The largest solid rocket motors ever built were Aerojet's three {{convert|260|in|m|adj=on|sigfig=3|order=flip|sp=us}} monolithic solid motors cast in Florida.<ref>{{Cite journal|url = https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20000033816.pdf|title = The 260 - The Largest Solid Rocket Motor Ever Tested|date = June 1999|access-date = July 24, 2014|website = nasa.gov}}</ref> Motors 260 SL-1 and SL-2 were {{convert|261|in|m|sigfig=3|order=flip|sp=us}} in diameter, {{convert|80|ft|8|in|m|order=flip|sp=us}} long, weighed {{convert|1,858,300|lb|kg|order=flip}}, and had a maximum thrust of {{convert|3.5|e6lbf|MN lbf|order=out|abbr=on}}. Burn duration was two minutes. The nozzle throat was large enough to walk through standing up. The motor was capable of serving as a 1-to-1 replacement for the 8-engine [[Saturn I]] liquid-propellant first stage but was never used as such. Motor 260 SL-3 was of similar length and weight but had a maximum thrust of {{convert|5.4|e6lbf|MN lbf|order=out|abbr=on}} and a shorter duration. | ||
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The following are chosen or solved simultaneously. The results are exact dimensions for grain, nozzle, and case geometries: | The following are chosen or solved simultaneously. The results are exact dimensions for grain, nozzle, and case geometries: | ||
* The grain burns at a predictable rate, given its surface area and chamber pressure.{{citation needed|date=June 2020}} | * The grain burns at a predictable rate, given its surface area and chamber pressure.{{citation needed|date=June 2020}}<ref>{{Cite book |last1=Kosanke |first1=K. L. |url=https://books.google.com/books?id=AK8zAwAAQBAJ&dq=the+grain+burns+at+a+predictable+rate%2C+given+its+surface+area+and+chamber+pressure&pg=PA888 |title=Encyclopedic Dictionary of Pyrotechnics: (and Related Subjects) |last2=Sturman |first2=Barry T. |last3=Winokur |first3=Robert M. |last4=Kosanke |first4=B. J. |date=October 2012 |publisher=Journal of Pyrotechnics |isbn=978-1-889526-21-8 |language=en}}</ref> | ||
* The chamber pressure is determined by the nozzle throat diameter and grain burn rate. | * The chamber pressure is determined by the nozzle throat diameter and grain burn rate. | ||
* Allowable chamber pressure is a function of casing design. | * Allowable chamber pressure is a function of casing design. | ||
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The casing must be designed to withstand the pressure and resulting stresses of the rocket motor, possibly at elevated temperature. For design, the casing is considered a [[pressure vessel]]. | The casing must be designed to withstand the pressure and resulting stresses of the rocket motor, possibly at elevated temperature. For design, the casing is considered a [[pressure vessel]]. | ||
To protect the casing from corrosive hot gases, a sacrificial thermal liner on the inside of the casing is often implemented, which [[ | To protect the casing from corrosive hot gases, a sacrificial thermal liner on the inside of the casing is often implemented, which [[Ablation|ablates]] to prolong the life of the motor casing. | ||
==Nozzle== | ==Nozzle== | ||
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A [[De Laval nozzle|convergent-divergent]] design accelerates the exhaust gas out of the nozzle to produce thrust. The nozzle must be constructed from a material that can withstand the heat of the combustion gas flow. Often, heat-resistant carbon-based materials are used, such as amorphous [[graphite]] or [[carbon-carbon]]. | A [[De Laval nozzle|convergent-divergent]] design accelerates the exhaust gas out of the nozzle to produce thrust. The nozzle must be constructed from a material that can withstand the heat of the combustion gas flow. Often, heat-resistant carbon-based materials are used, such as amorphous [[graphite]] or [[carbon-carbon]]. | ||
Some designs include directional control of the exhaust. This can be accomplished by gimballing the nozzle, as in the Space Shuttle SRBs, by the use of jet vanes in the exhaust as in the [[V-2]] rocket, or by liquid injection thrust vectoring (LITV). | Some designs include directional control of the exhaust. This can be accomplished by gimballing the nozzle, as in the Space Shuttle SRBs, by the use of jet vanes in the exhaust as in the [[V-2]] rocket, or by liquid injection thrust vectoring (LITV). | ||
LITV consists of injecting a liquid into the exhaust stream after the nozzle throat. The liquid then vaporizes, and in most cases chemically reacts, adding mass flow to one side of the exhaust stream and thus providing a control moment. For example, the [[Titan III]]C solid boosters injected [[nitrogen tetroxide]] for LITV; the tanks can be seen on the sides of the rocket between the main center stage and the boosters.<ref name="S">{{cite book | author=Sutton, George P. | title=Rocket Propulsion Elements | edition=7th | publisher=Wiley-Interscience | date=2000 | isbn=0-471-32642-9}}</ref> | LITV consists of injecting a liquid into the exhaust stream after the nozzle throat. The liquid then vaporizes, and in most cases chemically reacts, adding mass flow to one side of the exhaust stream and thus providing a control moment. For example, the [[Titan III]]C solid boosters injected [[nitrogen tetroxide]] for LITV; the tanks can be seen on the sides of the rocket between the main center stage and the boosters.<ref name="S">{{cite book | author=Sutton, George P. | title=Rocket Propulsion Elements | edition=7th | publisher=Wiley-Interscience | date=2000 | isbn=0-471-32642-9}}</ref> | ||
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|archive-date=30 July 2018 | |archive-date=30 July 2018 | ||
|url-status=dead | |url-status=dead | ||
}}</ref> This compares to {{cvt|339.3|isp}} for RP1/LOX (RD-180)<ref>http://www.pw.utc.com/Products/Pratt+%26+Whitney+Rocketdyne/Propulsion+Solutions/Space{{dead link|date=November 2017 |bot=InternetArchiveBot |fix-attempted=yes }}</ref> and {{cvt|452.3|isp}} for LH<sub>2</sub>/LOX (Block II [[RS-25]])<ref>{{cite web |url=http://www.pw.utc.com/Products/Pratt+%26+Whitney+Rocketdyne |title=Archived copy |access-date=2014-01-07 |url-status=dead |archive-url=https://web.archive.org/web/20110426223005/http://www.pw.utc.com/Products/Pratt+%2526+Whitney+Rocketdyne |archive-date=2011-04-26 }}</ref> bipropellant engines. Upper stage specific impulses are somewhat greater: as much as {{cvt|303.8|isp}} for APCP (Orbus 6E),<ref name="spaceandtech.com">{{cite web |url=http://www.spaceandtech.com/spacedata/elvs/titan4b_specs.shtml |title=Archived copy |access-date=2014-02-09 |url-status=dead |archive-url=https://web.archive.org/web/20130719221015/http://www.spaceandtech.com/spacedata/elvs/titan4b_specs.shtml |archive-date=2013-07-19 }}</ref> {{cvt|359|isp}} for RP1/LOX (RD-0124)<ref>http://www.russianspaceweb.com/engines/rd0124.htm</ref> and {{cvt|465.5|isp}} for LH<sub>2</sub>/LOX (RL10B-2).<ref>{{cite web |url=http://www.pw.utc.com/StaticFiles/Pratt%20.../Products/.../pwr_rl10b-2.pdf |title=RL10B-2 brochure|publisher=Pratt & Whitney Rocketdyne|year=2009 |access-date=2018-08-25 |url-status=dead |archive-url=https://web.archive.org/web/20120326211303/http://www.pw.utc.com/products/pwr/assets/pwr_rl10b-2.pdf |archive-date=2012-03-26 }}</ref> Propellant fractions are usually somewhat higher for (non-segmented) solid propellant first stages than for upper stages. The {{convert|117,000|lb|kg|order=flip|adj=on}} Castor 120 first stage has a propellant mass fraction of 92.23% while the {{convert|31,000|lb|kg|order=flip|adj=on}} Castor 30 upper stage developed for Orbital Science's Taurus II COTS(Commercial Off The Shelf) (International Space Station resupply) launch vehicle has a 91.3% propellant fraction with 2.9% graphite epoxy motor casing, 2.4% nozzle, igniter and thrust vector actuator, and 3.4% non-motor hardware including such things as payload mount, interstage adapter, cable raceway, instrumentation, etc. Castor 120 and Castor 30 are {{convert|93|and|92|in|m|2|order=flip|sp=us}} in diameter, respectively, and serve as stages on the Athena IC and IIC commercial launch vehicles. A four-stage Athena II using Castor 120s as both first and second stages became the first commercially developed launch vehicle to launch a lunar probe (''[[Lunar Prospector]]'') in 1998. | }}</ref> This compares to {{cvt|339.3|isp}} for RP1/LOX (RD-180)<ref>http://www.pw.utc.com/Products/Pratt+%26+Whitney+Rocketdyne/Propulsion+Solutions/Space{{dead link|date=November 2017 |bot=InternetArchiveBot |fix-attempted=yes }}</ref> and {{cvt|452.3|isp}} for LH<sub>2</sub>/LOX (Block II [[RS-25]])<ref>{{cite web |url=http://www.pw.utc.com/Products/Pratt+%26+Whitney+Rocketdyne |title=Archived copy |access-date=2014-01-07 |url-status=dead |archive-url=https://web.archive.org/web/20110426223005/http://www.pw.utc.com/Products/Pratt+%2526+Whitney+Rocketdyne |archive-date=2011-04-26 }}</ref> bipropellant engines. Upper stage specific impulses are somewhat greater: as much as {{cvt|303.8|isp}} for APCP (Orbus 6E),<ref name="spaceandtech.com">{{cite web |url=http://www.spaceandtech.com/spacedata/elvs/titan4b_specs.shtml |title=Archived copy |access-date=2014-02-09 |url-status=dead |archive-url=https://web.archive.org/web/20130719221015/http://www.spaceandtech.com/spacedata/elvs/titan4b_specs.shtml |archive-date=2013-07-19 }}</ref> {{cvt|359|isp}} for RP1/LOX (RD-0124)<ref>http://www.russianspaceweb.com/engines/rd0124.htm {{Dead link|date=February 2022}}</ref> and {{cvt|465.5|isp}} for LH<sub>2</sub>/LOX (RL10B-2).<ref>{{cite web |url=http://www.pw.utc.com/StaticFiles/Pratt%20.../Products/.../pwr_rl10b-2.pdf |title=RL10B-2 brochure|publisher=Pratt & Whitney Rocketdyne|year=2009 |access-date=2018-08-25 |url-status=dead |archive-url=https://web.archive.org/web/20120326211303/http://www.pw.utc.com/products/pwr/assets/pwr_rl10b-2.pdf |archive-date=2012-03-26 }}</ref> Propellant fractions are usually somewhat higher for (non-segmented) solid propellant first stages than for upper stages. The {{convert|117,000|lb|kg|order=flip|adj=on}} Castor 120 first stage has a propellant mass fraction of 92.23% while the {{convert|31,000|lb|kg|order=flip|adj=on}} Castor 30 upper stage developed for Orbital Science's Taurus II COTS(Commercial Off The Shelf) (International Space Station resupply) launch vehicle has a 91.3% propellant fraction with 2.9% graphite epoxy motor casing, 2.4% nozzle, igniter and thrust vector actuator, and 3.4% non-motor hardware including such things as payload mount, interstage adapter, cable raceway, instrumentation, etc. Castor 120 and Castor 30 are {{convert|93|and|92|in|m|2|order=flip|sp=us}} in diameter, respectively, and serve as stages on the Athena IC and IIC commercial launch vehicles. A four-stage Athena II using Castor 120s as both first and second stages became the first commercially developed launch vehicle to launch a lunar probe (''[[Lunar Prospector]]'') in 1998. | ||
Solid rockets can provide high thrust for relatively low cost. For this reason, solids have been used as initial stages in rockets (for example the [[Space Shuttle]]), while reserving high specific impulse engines, especially less massive hydrogen-fueled engines, for higher stages. In addition, solid rockets have a long history as the final boost stage for satellites due to their simplicity, reliability, compactness and reasonably high [[Propellant mass fraction|mass fraction]].<ref>[http://www.astronautix.com/props/solid.htm Solid<!-- Bot generated title -->] {{webarchive|url=https://web.archive.org/web/20020105224630/http://www.astronautix.com/props/solid.htm |date=2002-01-05 }}</ref> A spin-stabilized solid rocket motor is sometimes added when extra velocity is required, such as for a mission to a comet or the outer solar system, because a spinner does not require a guidance system (on the newly added stage). Thiokol's extensive family of mostly titanium-cased ''Star'' space motors has been widely used, especially on Delta launch vehicles and as spin-stabilized upper stages to launch satellites from the cargo bay of the Space Shuttle. ''Star'' motors have propellant fractions as high as 94.6% but add-on structures and equipment reduce the operating mass fraction by 2% or more. | Solid rockets can provide high thrust for relatively low cost. For this reason, solids have been used as initial stages in rockets (for example the [[Space Shuttle]]), while reserving high specific impulse engines, especially less massive hydrogen-fueled engines, for higher stages. In addition, solid rockets have a long history as the final boost stage for satellites due to their simplicity, reliability, compactness and reasonably high [[Propellant mass fraction|mass fraction]].<ref>[http://www.astronautix.com/props/solid.htm Solid<!-- Bot generated title -->] {{webarchive|url=https://web.archive.org/web/20020105224630/http://www.astronautix.com/props/solid.htm |date=2002-01-05 }}</ref> A spin-stabilized solid rocket motor is sometimes added when extra velocity is required, such as for a mission to a comet or the outer solar system, because a spinner does not require a guidance system (on the newly added stage). Thiokol's extensive family of mostly titanium-cased ''Star'' space motors has been widely used, especially on Delta launch vehicles and as spin-stabilized upper stages to launch satellites from the cargo bay of the Space Shuttle. ''Star'' motors have propellant fractions as high as 94.6% but add-on structures and equipment reduce the operating mass fraction by 2% or more. | ||
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A powdered oxidizer and powdered metal fuel are intimately mixed and immobilized with a rubbery binder (that also acts as a fuel). Composite propellants are often either [[ammonium nitrate]]-based (ANCP) or [[ammonium perchlorate]]-based (APCP). Ammonium nitrate composite propellant often uses [[magnesium]] and/or [[aluminium]] as fuel and delivers medium performance (I<sub>sp</sub> of about 210 s) whereas [[ammonium perchlorate composite propellant]] often uses aluminium fuel and delivers high performance (vacuum I<sub>sp</sub> up to 296 s with a single piece nozzle or 304 s with a high area ratio telescoping nozzle).<ref name="spaceandtech.com"/> Aluminium is used as fuel because it has a reasonable specific energy density, a high volumetric energy density, and is difficult to ignite accidentally. Composite propellants are cast, and retain their shape after the rubber binder, such as [[Hydroxyl-terminated polybutadiene]] (HTPB), [[cross-links]] (solidifies) with the aid of a curative additive. Because of its high performance, moderate ease of manufacturing, and moderate cost, APCP finds widespread use in space rockets, military rockets, hobby and amateur rockets, whereas cheaper and less efficient ANCP finds use in amateur rocketry and [[gas generator]]s. [[Ammonium dinitramide]], NH<sub>4</sub>N(NO<sub>2</sub>)<sub>2</sub>, is being considered as a 1-to-1 chlorine-free substitute for ammonium perchlorate in composite propellants. Unlike ammonium nitrate, ADN can be substituted for AP without a loss in motor performance. | A powdered oxidizer and powdered metal fuel are intimately mixed and immobilized with a rubbery binder (that also acts as a fuel). Composite propellants are often either [[ammonium nitrate]]-based (ANCP) or [[ammonium perchlorate]]-based (APCP). Ammonium nitrate composite propellant often uses [[magnesium]] and/or [[aluminium]] as fuel and delivers medium performance (I<sub>sp</sub> of about 210 s) whereas [[ammonium perchlorate composite propellant]] often uses aluminium fuel and delivers high performance (vacuum I<sub>sp</sub> up to 296 s with a single piece nozzle or 304 s with a high area ratio telescoping nozzle).<ref name="spaceandtech.com"/> Aluminium is used as fuel because it has a reasonable specific energy density, a high volumetric energy density, and is difficult to ignite accidentally. Composite propellants are cast, and retain their shape after the rubber binder, such as [[Hydroxyl-terminated polybutadiene]] (HTPB), [[cross-links]] (solidifies) with the aid of a curative additive. Because of its high performance, moderate ease of manufacturing, and moderate cost, APCP finds widespread use in space rockets, military rockets, hobby and amateur rockets, whereas cheaper and less efficient ANCP finds use in amateur rocketry and [[gas generator]]s. [[Ammonium dinitramide]], NH<sub>4</sub>N(NO<sub>2</sub>)<sub>2</sub>, is being considered as a 1-to-1 chlorine-free substitute for ammonium perchlorate in composite propellants. Unlike ammonium nitrate, ADN can be substituted for AP without a loss in motor performance. | ||
Polyurethane-bound aluminium-APCP solid fuel was used in the submarine launched [[Polaris missile]]s.<ref>{{Cite web|url=https://fas.org/nuke/guide/usa/slbm/a-1.htm|title = Polaris A1 - United States Nuclear Forces}}</ref> APCP used in the [[Space Shuttle Solid Rocket Boosters|space shuttle Solid Rocket Boosters]] consisted of ammonium perchlorate (oxidizer, 69.6% by weight), aluminium (fuel, 16%), iron oxide (a catalyst, 0.4%), polybutadiene acrylonitrile (PBAN) polymer (a non-urethane rubber binder that held the mixture together and acted as secondary fuel, 12.04%), and an epoxy [[Curing (chemistry)|curing]] agent (1.96%).<ref name="sts-newsref-srb">{{cite web | url = http://science.ksc.nasa.gov/shuttle/technology/sts-newsref/srb.html | title = Shuttle Solid Rocket Boosters | publisher = NASA }}</ref><ref name="returntoflight-system-SRB">{{cite web | url = http://www.nasa.gov/returntoflight/system/system_SRB.html | title = Solid Rocket Boosters | publisher = NASA}}</ref> It developed a specific impulse of 242 seconds (2.37 km/s) at sea level or 268 seconds (2.63 km/s) in a vacuum. The 2005-2009 [[Constellation Program]] was to use a similar PBAN-bound APCP.<ref>{{cite news |title=NASA Tests Engine With an Uncertain Future |url=https://www.nytimes.com/2010/08/31/science/space/31rocket.html?hpw |work=[[New York Times]] |date=August 30, 2010 |access-date=2010-08-31 | first=Kenneth | last=Chang}}</ref> | Polyurethane-bound aluminium-APCP solid fuel was used in the submarine launched [[Polaris missile]]s.<ref>{{Cite web|url=https://fas.org/nuke/guide/usa/slbm/a-1.htm|title = Polaris A1 - United States Nuclear Forces}}</ref> APCP used in the [[Space Shuttle Solid Rocket Boosters|space shuttle Solid Rocket Boosters]] consisted of ammonium perchlorate (oxidizer, 69.6% by weight), aluminium (fuel, 16%), iron oxide (a catalyst, 0.4%), polybutadiene acrylonitrile (PBAN) polymer (a non-urethane rubber binder that held the mixture together and acted as secondary fuel, 12.04%), and an epoxy [[Curing (chemistry)|curing]] agent (1.96%).<ref name="sts-newsref-srb">{{cite web | url = http://science.ksc.nasa.gov/shuttle/technology/sts-newsref/srb.html | title = Shuttle Solid Rocket Boosters | publisher = NASA | access-date = 2015-10-02 | archive-date = 2019-04-30 | archive-url = https://web.archive.org/web/20190430095734/https://science.ksc.nasa.gov/shuttle/technology/sts-newsref/srb.html | url-status = dead }}</ref><ref name="returntoflight-system-SRB">{{cite web | url = http://www.nasa.gov/returntoflight/system/system_SRB.html | title = Solid Rocket Boosters | publisher = NASA}}</ref> It developed a specific impulse of 242 seconds (2.37 km/s) at sea level or 268 seconds (2.63 km/s) in a vacuum. The 2005-2009 [[Constellation Program]] was to use a similar PBAN-bound APCP.<ref>{{cite news |title=NASA Tests Engine With an Uncertain Future |url=https://www.nytimes.com/2010/08/31/science/space/31rocket.html?hpw |work=[[New York Times]] |date=August 30, 2010 |access-date=2010-08-31 | first=Kenneth | last=Chang}}</ref> | ||
In 2009, a group succeeded in creating a propellant of [[water]] and nanoaluminium ([[ALICE (propellant)|ALICE]]). | In 2009, a group succeeded in creating a propellant of [[water]] and nanoaluminium ([[ALICE (propellant)|ALICE]]). | ||