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Liquid fuel cryogenic rocket engine, typically used on rocket upper stages From Wikipedia, the free encyclopedia
The RL10 is a liquid-fuel cryogenic rocket engine built in the United States by Aerojet Rocketdyne that burns cryogenic liquid hydrogen and liquid oxygen propellants. Modern versions produce up to 110 kN (24,729 lbf) of thrust per engine in vacuum. Three RL10 versions are in production for the Centaur upper stage of the Atlas V and the DCSS of the Delta IV. Three more versions are in development for the Exploration Upper Stage of the Space Launch System and the Centaur V of the Vulcan rocket.[2]
 | This article needs to be updated. The reason given is: Info on improvement projects is dated. (June 2020) |
.jpg) An RL10A-4 engine in London's Science Museum | |
| Country of origin | United States |
|---|---|
| First flight | 1962 (RL10A-1) |
| Manufacturer | Aerojet Rocketdyne |
| Application | Upper stage engine |
| Associated LV | Atlas · Delta III · Delta IV · Saturn I · SLS · Titan IIIE · Titan IV · Vulcan Centaur Canceled: DC-X · OmegA · Space Shuttle |
| Status | In production |
| Liquid-fuel engine | |
| Propellant | LOX / LH2 |
| Mixture ratio | 5.88:1 |
| Cycle | Expander cycle |
| Configuration | |
| Nozzle ratio | 84:1 or 280:1 |
| Performance | |
| Thrust, vacuum | 110.1 kN (24,800 lbf) |
| Specific impulse, vacuum | 465.5 s (4.565 km/s) |
| Dimensions | |
| Length | 4.15 m (13.6 ft) w/ nozzle extended |
| Diameter | 2.15 m (7 ft 1 in) |
| Dry mass | 301 kg (664 lb) |
| Used in | |
| Centaur, DCSS, S-IV | |
| References | |
| References | [1] |
| Notes | Performance values and dimensions are for RL10B-2. |
The expander cycle that the engine uses drives the turbopump with waste heat absorbed by the engine combustion chamber, throat, and nozzle. This, combined with the hydrogen fuel, leads to very high specific impulses (Isp) in the range of 373 to 470 s (3.66–4.61 km/s) in a vacuum. Mass ranges from 131 to 317 kg (289–699 lb) depending on the version of the engine.[3][4]
The RL10 was the first liquid hydrogen rocket engine to be built in the United States, with development of the engine by Marshall Space Flight Center and Pratt & Whitney beginning in the 1950s. The RL10 was originally developed as a throttleable engine for the USAF Lunex lunar lander.[5]
The RL10 was first tested on the ground in 1959, at Pratt & Whitney's Florida Research and Development Center in West Palm Beach, Florida.[6][7] The first successful flight took place on November 27, 1963.[8][9] For that launch, two RL10A-3 engines powered the Centaur upper stage of an Atlas launch vehicle. The launch was used to conduct a heavily instrumented performance and structural integrity test of the vehicle.[10]
Multiple versions of this engine have been flown. The S-IV of the Saturn I used a cluster of six RL10A-3S, a version which was modified for installation on the Saturn[11] and the Titan program included Centaur D-1T upper stages powered by two RL10A-3-3 Engines.[11][12]
Four modified RL10A-5 engines were used in the McDonnell Douglas DC-X.[13]
A flaw in the brazing of an RL10B-2 combustion chamber was identified as the cause of failure for the 4 May 1999 Delta III launch carrying the Orion-3 communications satellite.[14]
The DIRECT version 3.0 proposal to replace Ares I and Ares V with a family of rockets sharing a common core stage recommended the RL10 for the second stage of the J-246 and J-247 launch vehicles.[15] Up to seven RL10 engines would have been used in the proposed Jupiter Upper Stage, serving an equivalent role to the Space Launch System Exploration Upper Stage.
In the early 2000s, NASA contracted with Pratt & Whitney Rocketdyne to develop the Common Extensible Cryogenic Engine (CECE) demonstrator. CECE was intended to lead to RL10 engines capable of deep throttling.[16] In 2007, its operability (with some "chugging") was demonstrated at 11:1 throttle ratios.[17] In 2009, NASA reported successfully throttling from 104 percent thrust to eight percent thrust, a record for an expander cycle engine of this type. Chugging was eliminated by injector and propellant feed system modifications that control the pressure, temperature and flow of propellants.[18] In 2010, the throttling range was expanded further to a 17.6:1 ratio, throttling from 104% to 5.9% power.[19]
In 2012 NASA joined with the US Air Force (USAF) to study next-generation upper stage propulsion, formalizing the agencies' joint interests in a new upper stage engine to replace the Aerojet Rocketdyne RL10.
"We know the list price on an RL10. If you look at cost over time, a very large portion of the unit cost of the EELVs is attributable to the propulsion systems, and the RL10 is a very old engine, and there's a lot of craftwork associated with its manufacture. ... That's what this study will figure out, is it worthwhile to build an RL10 replacement?"
— Dale Thomas, Associated Director Technical, Marshall Space Flight Center[20]
From the study, NASA hoped to find a less expensive RL10-class engine for the upper stage of the Space Launch System (SLS).[20][21]
USAF hoped to replace the Rocketdyne RL10 engines used on the upper stages of the Lockheed Martin Atlas V and the Boeing Delta IV Evolved Expendable Launch Vehicles (EELV) that were the primary methods of putting US government satellites into space.[20] A related requirements study was conducted at the same time under the Affordable Upper Stage Engine Program (AUSEP).[21]
The RL10 has evolved over the years. The RL10B-2 that was used on the DCSS had improved performance, an extendable carbon-carbon nozzle, electro-mechanical gimbaling for reduced weight and increased reliability, and a specific impulse of 465.5 seconds (4.565 km/s).[22][23]
As of 2016, Aerojet Rocketdyne was working toward incorporating additive manufacturing into the RL10 construction process. The company conducted full-scale, hot-fire tests on an engine with a printed main injector in March 2016,[24] and on an engine with a printed thrust chamber assembly in April 2017.[25]
As of 2009[update], an enhanced version of the RL10 was proposed to power the Advanced Cryogenic Evolved Stage (ACES), a long-duration, low-boiloff extension of existing ULA Centaur and Delta Cryogenic Second Stage (DCSS) technology for the Vulcan launch vehicle.[34] Long-duration ACES technology is intended to support geosynchronous, cislunar, and interplanetary missions. Another possible application is as in-space propellant depots in LEO or at L2 that could be used as way-stations for other rockets to stop and refuel on the way to beyond-LEO or interplanetary missions. Cleanup of space debris was also proposed.[35]
| Version | Status | First flight | Dry mass | Thrust | Isp (ve), vac. | Length | Nozzle diameter | T:W | O:F | Expansion ratio | Chamber pressure | Burn time | Associated stage | Notes |
|---|---|---|---|---|---|---|---|---|---|---|---|---|---|---|
| RL10A-1 | Retired | 1962 | 131 kg (289 lb) | 67 kN (15,000 lbf) | 425 s (4.17 km/s) | 1.73 m (5 ft 8 in) | 1.53 m (5 ft 0 in) | 52:1 | 5:1 | 40:1 | 20.7 bar (2,070 kPa) | 430 s | Centaur A | Prototype [11][26][36][37] |
| RL10A-3C | Retired | 1963 | 131 kg (289 lb) | 65.6 kN (14,700 lbf) | 444 s (4.35 km/s) | 2.49 m (8 ft 2 in) | 1.53 m (5 ft 0 in) | 51:1 | 5:1 | 57:1 | 32.75 bar (3,275 kPa) | 470 s | Centaur B/C/D/E | [38] |
| RL10A-3S | Retired | 1964 | 134 kg (296 lb) | 67 kN (15,000 lbf) | 427 s (4.19 km/s) | 1.73 m (5 ft 8 in) | 51:1 | 5:1 | 40:1 | 20.7 bar (2,070 kPa) | S-IV | [11][8] | ||
| RL10A-4 | Retired | 1992 | 168 kg (370 lb) | 92.5 kN (20,800 lbf) | 449 s (4.40 km/s) | 2.29 m (7 ft 6 in) | 1.17 m (3 ft 10 in) | 56:1 | 5.5:1 | 84:1 | 39.8 bar (3,980 kPa) | 392 s | Centaur IIA | [11][39] |
| RL10A-5 | Retired | 1993 | 143 kg (315 lb) | 64.7 kN (14,500 lbf) | 373 s (3.66 km/s) | 1.07 m (3 ft 6 in) | 1.02 m (3 ft 4 in) | 46:1 | 6:1 | 4:1 | 39.8 bar (3,980 kPa) | 127 s | DC-X | [11][40] |
| RL10B-2 | Retired | 1998 | 301 kg (664 lb) | 110.1 kN (24,750 lbf) | 465.5 s (4.565 km/s) | Stowed: 2.2 m (7 ft 2.5 in) Deployed: 4.15 m (13 ft 7.5 in) |
2.15 m (7 ft 0.5 in) | 40:1 | 5.88:1 | 280:1 | 44.12 bar (4,412 kPa) | 5m: 1,125 s 4m: 700 s |
DCSS, ICPS | Succeeded by RL10C-2.[1][41] |
| RL10A-4-1 | Retired | 2000 | 167 kg (368 lb) | 99.1 kN (22,300 lbf) | 451 s (4.42 km/s) | 1.78 m (5 ft 10 in) | 1.53 m (5 ft 0 in) | 61:1 | 84:1 | 42 bar (4,200 kPa) | 740 s | Centaur IIIA | [11][42] | |
| RL10A-4-2 | Active | 2002 | 170 kg (370 lb) | 99 kN (22,300 lbf) | 451 s (4.42 km/s) | 2.29 m (7 ft 6 in) | 1.17 m (3 ft 10 in) | 61:1 | 84:1 | 42 bar (4,200 kPa) | 740 s | Centaur IIIB Centaur SEC Centaur DEC |
Used for Starliner launches.[11][43][44] | |
| RL10B-X | Cancelled | — | 317 kg (699 lb) | 93.4 kN (21,000 lbf) | 470 s (4.6 km/s) | 1.53 m (5 ft 0 in) | 30:1 | 250:1 | 408 s | Centaur B-X | [45] | |||
| CECE | Demonstrator project | — | 160 kg (350 lb) | 67 kN (15,000 lbf), throttle to 5–10% | >445 s (4.36 km/s) | 1.53 m (5 ft 0 in) | 43:1 | [46][47] | ||||||
| RL10C-1 | Retired | 2014 | 190 kg (420 lb) | 101.5 kN (22,820 lbf) | 449.7 s (4.410 km/s) | 2.18 m (7 ft 2 in) | 1.45 m (4 ft 9 in) | 57:1 | 5.5:1 | 130:1 | Centaur SEC Centaur DEC |
Succeeded by RL-10C-1-1.[48][49][50][44] | ||
| RL10C-1-1 | Active | 2021 | 188 kg (415 lb) | 105.98 kN (23,825 lbf) | 453.8 s (4.450 km/s) | 2.46 m (8 ft 0.7 in) | 1.57 m (5 ft 2 in) | 57:1 | 5.5:1 | 155:1 | Centaur SEC Centaur V |
Current standard engine for Atlas V and Vulcan Centaur.[11][2] | ||
| RL10C-2-1 | Retired | 2022 | 301 kg (664 lb) | 110.1 kN (24,750 lbf) | 465.5 s (4.565 km/s) | Stowed: 2.2 m (7 ft 2.5 in) Deployed: 4.15 m (13 ft 7.5 in) |
2.15 m (7 ft 0.5 in) | 37:1 | 5.88:1 | 280:1 | DCSS | [51][52] | ||
| RL10C-2 | Delivered, not yet flown | 2025 (expected) | 110.1 kN (24,750 lbf) | 465.5 s (4.565 km/s) | 4.15 m (13 ft 7 in) | 2.15 m (7 ft 1 in) | 37:1 | 5.88:1 | 280:1 | ICPS | Conversion of C-3[53] | |||
| RL10C-3 | Delivered, not yet flown | 2028 (expected) | 230 kg (508 lb) | 108.3 kN (24,340 lbf) | 460.1 s (4.512 km/s) | 3.16 m (10 ft 4.3 in) | 1.85 m (6 ft 1 in) | 48:1 | 5.7:1 | 215:1 | Exploration Upper Stage | [11][2][53] | ||
| RL10C-5-1 | Cancelled | — | 188 kg (414 lb) | 106 kN (23,825 lbF) | 453.8 s (4.450 km/s) | 2.46 m (8 ft 0.7 in) | 1.57 m (4 ft 9 in) | 57:1 | 5.5:1 | OmegA | [2][33] | |||
| RL10C-X | In development | 2025 (expected) | 231 kg (509 lb) | 107.29 kN (24,120 lbF) | 460.9 s (4.520 km/s) | 3.31 m (130.4 in) | 1.87 m (73.7 in) | 47.29:1 | 5.5:1 | Centaur V | Additive manufacturing [54][55] |
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