Jump to content

TR-201: Difference between revisions

From Wikipedia, the free encyclopedia
Browse history interactively
← Previous edit
Content deleted Content added
VisualWikitext
No edit summary
Cleanup infobox
 
(25 intermediate revisions by 16 users not shown)
Line 1: Line 1:
{{Short description|Rocket engine}}
{{Infobox rocket engine
{{Infobox rocket engine
|country_of_origin = [[United States]]
| country_of_origin = [[United States]]
|image =
| image =
|image_size =
| image_size =
|caption =
| caption =
|name = TR-201
| name = TR-201
|date = 1972–1988
| date = 1972–1988
|manufacturer = [[TRW Inc.|TRW]]
| manufacturer = [[TRW Inc.|TRW]]
|purpose = Upper stage/[[Spacecraft]] propulsion
| purpose = Upper stage
⚫
| status = Retired
⚫
|predecessor = [[Descent Propulsion System|LMDE]]
⚫
| cycle = [[Pressure-fed engine (rocket)|Pressure-fed engine]]
⚫
|status = Retired
⚫
| combustion_chamber = 1
⚫
|cycle = [[Pressure-fed engine (rocket)|Pressure-fed engine]]
| thrust(Vac) = {{cvt|41.9|kN}}
⚫
|combustion_chamber = 1
⚫
| specific_impulse_vacuum = {{cvt|301|isp}}
|thrust(Vac) = 41.90{{nbsp}}kN (9,419{{nbsp}}lbf)
⚫
| chamber_pressure = {{cvt|700|kPa|psi}}
⚫
|specific_impulse_vacuum = 301{{nbsp}}s (3,050{{nbsp}}N⋅s/kg)
⚫
| thrust_to_weight = 31.4
⚫
|chamber_pressure = 7.00{{nbsp}}bar
| diameter = {{cvt|1.38|m}}
⚫
|thrust_to_weight = 31.4
| length = {{cvt|2.27|m}}
|diameter = 1.38{{nbsp}}m (4.52{{nbsp}}ft)
| dry_weight = {{cvt|113|kg}}
|length = 2.27{{nbsp}}m (7.44{{nbsp}}ft)
⚫
| predecessor = [[Descent Propulsion System|LMDE]]
|dry_weight = 113{{nbsp}}kg (249{{nbsp}}lb)
| successor =
|predecessor = [[Descent Propulsion System|LMDE]]
| type = liquid
|successor =
⚫
| fuel = [[Aerozine 50]]
|type = Liquid
| oxidiser = {{chem2|N2O4|link=nitrogen tetroxide}}
⚫
|fuel = [[Aerozine 50]]
⚫
| used_in = [[Delta-P]], second stage of [[Delta (rocket family)]]
|oxidiser = [[N2O4|N{{sub|2}}O{{sub|4}}]]
⚫
|used_in = [[Delta-P]], second stage of [[Delta (rocket family)]]
}}
}}

The '''TR-201 or TR201''' is a [[hypergolic]] pressure-fed [[rocket engine]] used to propel the upper stage of the [[Delta (rocket family)|Delta rocket]], referred to as [[Delta-P]], from 1972 to 1988.
The '''TR-201 or TR201''' is a [[hypergolic]] pressure-fed [[rocket engine]] used to propel the upper stage of the [[Delta (rocket family)|Delta rocket]], referred to as [[Delta-P]], from 1972 to 1988.
The [[rocket engine]] uses [[Aerozine 50]] as fuel, and [[dinitrogen tetroxide|{{chem|N|2|O|4}}]] as oxidizer. It was developed in early 1970s by [[TRW Inc.|TRW]] as a derivative of the [[Descent Propulsion System|Lunar Module Descent Engine (LMDE)]]. This engine used a [[pintle injector]] first developed by TRW in late 1950s and received US Patent in 1972.<ref name="trwpintle">{{cite journal |url=http://smartdata.usbid.com/datasheets/usbid/2001/2001-q1/pintleenginepaperaiaafinal.pdf |last=Dressler|first=Gordan A.|author2=Bauer, J. Martin |year=2000|title=TRW Pintle Engine Heritage and Performance Characteristics |id=AIAA-2000-3871|publisher=[[AIAA]] |accessdate=4 June 2012 }}</ref> This injector technology and design is also used on SpaceX [[Merlin (rocket engine family)|Merlin]] engines.<ref>{{cite web |url=http://www.astronautix.com/engines/tr201.htm |title=TR-201 |publisher=Encyclopedia Astronautica |accessdate=4 June 2012 |deadurl=yes |archiveurl=https://web.archive.org/web/20080706190928/http://astronautix.com/engines/tr201.htm |archivedate=6 July 2008 |df= }}</ref>
The [[rocket engine]] uses [[Aerozine 50]] as fuel, and [[dinitrogen tetroxide|{{chem|N|2|O|4}}]] as oxidizer. It was developed in the early 1970s by [[TRW Inc.|TRW]] as a derivative of the [[Descent propulsion system|lunar module descent engine (LMDE)]]. This engine used a [[pintle injector]] first invented by Gerard W. Elverum Jr.<ref>{{Cite patent|country=US Patent|number=3,205,656|title=Variable thrust bipropellant rocket engine|status=|pubdate=|gdate=1963-02-25|invent1=Elverum Jr.|inventor1-first=Gerard W.|url=https://patents.google.com/patent/US3205656A/en}}</ref><ref>{{Cite patent|country=US Patent|number=3,699,772|title=Liquid propellant rocket engine coaxial injector|status=|pubdate=|gdate=1968-01-08|invent1=Elverum Jr.|inventor1-first=Gerard W.|url=https://patents.google.com/patent/US3699772A/en}}</ref><ref>{{Cite book|title=REMEMBERING THE GIANTS - Apollo Rocket Propulsion Development|publisher=NASA|pages=73–86}}</ref> and developed by TRW in the late 1950s and received US Patent in 1972.<ref name="trwpintle">{{cite conference|title=TRW Pintle Engine Heritage and Performance Characteristics|url=http://www.rocket-propulsion.info/resources/articles/TRW_PINTLE_ENGINE.pdf |conference=36th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit |date=July 2000 |archive-url=https://web.archive.org/web/20220331082846/http://www.rocket-propulsion.info/resources/articles/TRW_PINTLE_ENGINE.pdf |archive-date=31 March 2022 |url-status=live |doi=10.2514/6.2000-3871 |location=Redondo Beach, CA |publisher=TRW Inc. |last1=Dressler |first1=Gordon A. |last2=Bauer |first2=J. Martin |id=AIAA 2000-3871 }}</ref> This injector technology and design is also used on SpaceX [[Merlin (rocket engine family)|Merlin]] engines.<ref>{{cite web |url=http://www.astronautix.com/engines/tr201.htm |title=TR-201 |publisher=Encyclopedia Astronautica |access-date=4 June 2012 |url-status=dead |archive-url=https://web.archive.org/web/20080706190928/http://astronautix.com/engines/tr201.htm |archive-date=6 July 2008 }}</ref>


The thrust chamber was initially developed for the Apollo Lunar Module and was subsequently adopted for the Delta Expendable Launch Vehicle 2nd stage. The engine made 10 flights during the Apollo program and 77 during its Delta career between 1974-1988.
The thrust chamber was initially developed for the [[Apollo Lunar Module]] and was subsequently adopted for the Delta expendable launch vehicle 2nd stage. The engine made 10 flights during the Apollo program and 77 during its Delta career between 1974 and 1988.
The TRW TR-201 was re-configured as a fixed thrust version of the Lunar Module Descent Engine (LMDE) for Delta's stage 2.
The TRW TR-201 was re-configured as a fixed-thrust version of the LMDE for Delta's stage 2.
Multi start operation is adjustable up to 55.6&nbsp;kN and propellant throughput up to 7,711&nbsp;kg; and the engine can be adapted to optional expansion ratio nozzles. Development of the innovative thrust chamber and pintle design is credited to TRW Aerospace Engineer Dr. Peter Staudhammer.<ref>{{cite conference |url=https://arc.aiaa.org/doi/abs/10.2514/6.1967-521 |title=The Descent Engine for the Lunar Module |last1=Elverum |first1=Gordon |last2=Staudhammer |first2=Peter |others=J. Miller, A. Hoffman, and R. Rockow |publisher=AIAA |volume=67-521 |conference=AIAA 3rd Propulsion Joint Specialist Conference |date=17 July 1967}}</ref>
Multi-start operation is adjustable up to 55.6&nbsp;kN and propellant throughput up to 7,711&nbsp;kg; and the engine can be adapted to optional expansion ratio nozzles. Development of the innovative thrust chamber and pintle design is credited to TRW Aerospace Engineer Gerard W. Elverum Jr.<ref>{{Cite patent|country=US Patent|number=3,699,772|title=Liquid propellant rocket engine coaxial injector|status=|pubdate=|gdate=1968-01-08|invent1=Elverum Jr.|inventor1-first=Gerard W.|url=https://patents.google.com/patent/US3699772A/en}}</ref><ref>{{Cite patent|country=US Patent|number=3,205,656|title=Variable thrust bipropellant rocket engine|status=|pubdate=|gdate=1963-02-25|invent1=Elverum Jr.|inventor1-first=Gerard W.|url=https://patents.google.com/patent/US3205656A/en}}</ref>


The combustion chamber consists of an ablative-lined titanium alloy case to the 16:1 area ratio.
The combustion chamber consists of an ablative-lined [[titanium alloy]] case to the 16:1 area ratio.
Fabrication of the 6Al4V alloy titanium case was accomplished by machining the chamber portion and the exit cone portion from forgings and welding them into one unit at the throat centerline. The ablative liner is fabricated in two segments and installed from either end.
Fabrication of the [[Ti6Al4V]] alloy titanium case was accomplished by machining the chamber portion and the exit cone portion from forgings and welding them into one unit at the throat centerline. The ablative liner is fabricated in two segments and installed from either end.
The shape of the nozzle extension is such that the ablative liner is retained in the exit cone during transportation, launch and boost.
The shape of the nozzle extension is such that the ablative liner is retained in the exit cone during transportation, launch and boost.
During engine firing, thrust loads force the exit cone liner against the case.
During engine firing, thrust loads force the exit cone liner against the case.
The titanium head end assembly which contains the Pintle Injector and propellant valve subcomponents is attached with 36 A-286 steel ¼ inch bolts.
The titanium head end assembly which contains the Pintle Injector and propellant valve subcomponents is attached with 36 A-286 steel {{convert|1/4|in}} bolts.


In order to keep the maximum operating temperatures of the titanium case in the vicinity of 800{{nbsp}}°F, the ablative liner was designed as a composite material providing the maximum heat sink and minimum weight. The selected configuration consisted of a high density, erosion-resistant silica cloth/phenolic material surrounded by a lightweight needle-felted silica mat/phenolic insulation.
In order to keep the maximum operating temperatures of the titanium case in the vicinity of {{cvt|800|°F}}, the ablative liner was designed as a composite material providing the maximum heat sink and minimum weight. The selected configuration consisted of a high density, erosion-resistant silica cloth/phenolic material surrounded by a lightweight needle-felted silica mat/phenolic insulation.


The installed pintle injector, unique to TRW designed liquid propulsion systems, provides improved reliability and less costly method of fuel-oxidizer impingement in the thrust chamber than conventional coaxial distributed-element injectors typically used on liquid bipropellant rocket engines.
The installed pintle injector, unique to TRW-designed liquid-propulsion systems, provides improved reliability and less costly method of fuel–oxidizer impingement in the thrust chamber than conventional coaxial distributed-element injectors typically used on liquid bipropellant rocket engines.


==Specifications==
==Specifications==
* Number flown: 77 (Delta 2000 configuration)
* Number flown: 77 (Delta 2000 configuration)
* Dry mass: 300 pounds with Columbium ([[Niobium]]) nozzle extension installed
* Dry mass: {{convert|300|lb}} with columbium ([[niobium]]) [[nozzle extension]] installed
* Length: 51 inches - Gimbal attachment to nozzle tip (minus nozzle extension)
* Length: 51 inches – gimbal attachment to nozzle tip (minus nozzle extension)
* Maximum diameter: 34 inches (minus nozzle extension)
* Maximum diameter: {{convert|34|in}} (minus nozzle extension)
* Mounting: gimbal attachment above injector
* Mounting: gimbal attachment above injector
* Engine cycle: pressure fed (15.5 atm reservoir)
* Engine cycle: pressure fed (15.5 atm reservoir)
* Fuel: 50:50 N{{sub|2}}O{{sub|4}}/UDMH ([[Aerozine-50]]) at 8.92&nbsp;kg/s
* Fuel: 50:50 N{{sub|2}}O{{sub|4}}/UDMH ([[Aerozine 50]]) at 8.92&nbsp;kg/s
* Oxidizer: Dinitrogen tetroxide at 5.62&nbsp;kg/s
* Oxidizer: dinitrogen tetroxide at 5.62&nbsp;kg/s
* Oxidizer:fuel ratio: 1.60
* Oxidizer:fuel ratio: 1.60
* Thrust, vacuum: 42.923&nbsp;kN
* Thrust, vacuum: 42.923&nbsp;kN
Line 63: Line 64:
* Burn time: 500{{nbsp}}s for total of 5 starts; 10 × 350-s single burn
* Burn time: 500{{nbsp}}s for total of 5 starts; 10 × 350-s single burn


==Delta Usage==
==Delta usage==
The TR-201 engine was used as the second stage for 77 [[Delta (rocket family)|Delta]] launches between 1972 and 1988. The engine had a 100% reliability record during this 15 year operational period.<ref>{{cite web |url=http://www.astronautix.com/stages/deltap.htm |title=Delta P |publisher=Encyclopedia Astronautica |accessdate=4 June 2012 |deadurl=yes |archiveurl=https://web.archive.org/web/20120617125204/http://www.astronautix.com/stages/deltap.htm |archivedate=17 June 2012 |df= }}</ref>
The TR-201 engine was used as the second stage for 77 [[Delta (rocket family)|Delta]] launches between 1972 and 1988. The engine had a 100% reliability record during this 15 year operational period.<ref>{{cite web |url=http://www.astronautix.com/stages/deltap.htm |title=Delta P |publisher=Encyclopedia Astronautica |access-date=4 June 2012 |url-status=dead |archive-url=https://web.archive.org/web/20120617125204/http://www.astronautix.com/stages/deltap.htm |archive-date=17 June 2012 }}</ref>


== References ==
== References ==
Line 71: Line 72:
{{Rocket engines}}
{{Rocket engines}}
{{Orbital spacecraft rocket engines}}
{{Orbital spacecraft rocket engines}}

== External links ==


[[Category:Rocket engines using hypergolic propellant]]
[[Category:Rocket engines using hypergolic propellant]]
[[Category:Rocket engines using the pressure-fed cycle]]
[[Category:Rocket engines using the pressure-fed cycle]]
[[Category:Rocket engines of the United States]]

The claim that Dr. Peter Staudhammer is the inventor of the pintle throttle engine is inaccurate, and can be verified by the patent held by Gerald W. Elverum, Jr. The following is the link to the actual patent held by Mr. Elverum and the claim should be corrected to show the actual inventor of the pintle engine. (Use the link below to connect you to the U.S.P.O.)

http://patft.uspto.gov/netacgi/nph-Parser?Sect2=PTO1&Sect2=HITOFF&p=1&u=%2Fnetahtml%2FPTO%2Fsearch-bool.html&r=1&f=G&l=50&d=PALL&RefSrch=yes&Query=PN%2F3699772

While Dr. Staudhammer may have and most likely was involved with the program, an actual and accurate portrayal of his role in the project should be written and inserted into the Wikipedia article TR-201 (new section). Mr. Elverum's. Patent is evidence of ownership and authenticates him as being the "Soul inventor" of the pintle engine, and the article TR-201 (new section) should reflect the true and accurate facts in the article.

Furthermore, there have been several documentaries by various media outlets, Public Broadcasting Service, the Smithsonian and the likes that have interviewed, and produced media reflecting Mr. Elverum's invention, and the story behind the pintle engines design, operation and manufacturing.

Latest revision as of 23:53, 14 July 2024

TR-201
Country of originUnited States
Date1972–1988
ManufacturerTRW
ApplicationUpper stage
PredecessorLMDE
StatusRetired
Liquid-fuel engine
PropellantN2O4 / Aerozine 50
CyclePressure-fed engine
Configuration
Chamber1
Performance
Thrust, vacuum41.9 kN (9,400 lbf)
Thrust-to-weight ratio31.4
Chamber pressure700 kPa (100 psi)
Specific impulse, vacuum301 s (2.95 km/s)
Dimensions
Length2.27 m (7 ft 5 in)
Diameter1.38 m (4 ft 6 in)
Dry mass113 kg (249 lb)
Used in
Delta-P, second stage of Delta (rocket family)

The TR-201 or TR201 is a hypergolic pressure-fed rocket engine used to propel the upper stage of the Delta rocket, referred to as Delta-P, from 1972 to 1988. The rocket engine uses Aerozine 50 as fuel, and N
2O
4 as oxidizer. It was developed in the early 1970s by TRW as a derivative of the lunar module descent engine (LMDE). This engine used a pintle injector first invented by Gerard W. Elverum Jr.[1][2][3] and developed by TRW in the late 1950s and received US Patent in 1972.[4] This injector technology and design is also used on SpaceX Merlin engines.[5]

The thrust chamber was initially developed for the Apollo Lunar Module and was subsequently adopted for the Delta expendable launch vehicle 2nd stage. The engine made 10 flights during the Apollo program and 77 during its Delta career between 1974 and 1988. The TRW TR-201 was re-configured as a fixed-thrust version of the LMDE for Delta's stage 2. Multi-start operation is adjustable up to 55.6 kN and propellant throughput up to 7,711 kg; and the engine can be adapted to optional expansion ratio nozzles. Development of the innovative thrust chamber and pintle design is credited to TRW Aerospace Engineer Gerard W. Elverum Jr.[6][7]

The combustion chamber consists of an ablative-lined titanium alloy case to the 16:1 area ratio. Fabrication of the Ti6Al4V alloy titanium case was accomplished by machining the chamber portion and the exit cone portion from forgings and welding them into one unit at the throat centerline. The ablative liner is fabricated in two segments and installed from either end. The shape of the nozzle extension is such that the ablative liner is retained in the exit cone during transportation, launch and boost. During engine firing, thrust loads force the exit cone liner against the case. The titanium head end assembly which contains the Pintle Injector and propellant valve subcomponents is attached with 36 A-286 steel 1⁄4 inch (6.4 mm) bolts.

In order to keep the maximum operating temperatures of the titanium case in the vicinity of 800 °F (427 °C), the ablative liner was designed as a composite material providing the maximum heat sink and minimum weight. The selected configuration consisted of a high density, erosion-resistant silica cloth/phenolic material surrounded by a lightweight needle-felted silica mat/phenolic insulation.

The installed pintle injector, unique to TRW-designed liquid-propulsion systems, provides improved reliability and less costly method of fuel–oxidizer impingement in the thrust chamber than conventional coaxial distributed-element injectors typically used on liquid bipropellant rocket engines.

Specifications

[edit]
  • Number flown: 77 (Delta 2000 configuration)
  • Dry mass: 300 pounds (140 kg) with columbium (niobium) nozzle extension installed
  • Length: 51 inches – gimbal attachment to nozzle tip (minus nozzle extension)
  • Maximum diameter: 34 inches (860 mm) (minus nozzle extension)
  • Mounting: gimbal attachment above injector
  • Engine cycle: pressure fed (15.5 atm reservoir)
  • Fuel: 50:50 N2O4/UDMH (Aerozine 50) at 8.92 kg/s
  • Oxidizer: dinitrogen tetroxide at 5.62 kg/s
  • Oxidizer:fuel ratio: 1.60
  • Thrust, vacuum: 42.923 kN
  • Specific impulse, vacuum: 303 s
  • Expansion ratio: 16:1 without nozzle extension; 43:1 with nozzle extension
  • Cooling, upper thrust chamber: film
  • Cooling, lower thrust chamber: ablative quartz phenolic;
  • Cooling, nozzle extension: radiative
  • Chamber pressure: 7.1 atm
  • Ignition: hypergolic, started by 28 V electrical signal to on/off solenoid valves
  • Burn time: 500 s for total of 5 starts; 10 × 350-s single burn

Delta usage

[edit]

The TR-201 engine was used as the second stage for 77 Delta launches between 1972 and 1988. The engine had a 100% reliability record during this 15 year operational period.[8]

References

[edit]
  1. ^ US Patent 3,205,656, Elverum Jr., Gerard W., "Variable thrust bipropellant rocket engine", issued 1963-02-25 
  2. ^ US Patent 3,699,772, Elverum Jr., Gerard W., "Liquid propellant rocket engine coaxial injector", issued 1968-01-08 
  3. ^ REMEMBERING THE GIANTS - Apollo Rocket Propulsion Development. NASA. pp. 73–86.
  4. ^ Dressler, Gordon A.; Bauer, J. Martin (July 2000). TRW Pintle Engine Heritage and Performance Characteristics (PDF). 36th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit. Redondo Beach, CA: TRW Inc. doi:10.2514/6.2000-3871. AIAA 2000-3871. Archived (PDF) from the original on 31 March 2022.
  5. ^ "TR-201". Encyclopedia Astronautica. Archived from the original on 6 July 2008. Retrieved 4 June 2012.
  6. ^ US Patent 3,699,772, Elverum Jr., Gerard W., "Liquid propellant rocket engine coaxial injector", issued 1968-01-08 
  7. ^ US Patent 3,205,656, Elverum Jr., Gerard W., "Variable thrust bipropellant rocket engine", issued 1963-02-25 
  8. ^ "Delta P". Encyclopedia Astronautica. Archived from the original on 17 June 2012. Retrieved 4 June 2012.
Retrieved from "https://en.wikipedia.org/enwiki/w/index.php?title=TR-201&oldid=1234556344"