Cool the engines
Pratt & Whitney Rocketdyne, already busy upgrading its proven RS-68 engine on the Delta IV, may get even busier depending on which direction the Augustine Committee and the president decide to go with Project Constellation.
The current evolutionary roadmap for RS-68 goes to RS-68A, funded by DoD to improve performance for the Delta IV Heavy and eventually the entire Delta IV fleet. By upgrading the turbopumps for higher flowrates, PWR will put more injectors into the thrust chamber of RS-68A to improve thrust and specific impluse (raising from 409 sec to 414 sec in vacuum.)
NASA's current roadmap relied on a further upgrade of RS-68, the RS-68B model, for the Ares V super-booster. The NASA-funded upgrade would increase the engine's redundancy in certain systems (required by the current man-rating rules, and a focal point of the Delta IV crew-launch studies) and reducing the amount of gas that collects at the launch pad prior to ignition.
Yet the findings of the DIRECT team prior to unveiling DIRECT 3.0 may put a wrench in NASA's plans. Their studies (and apparently NASA's internal studies confirm this) indicate the current RS-68 engines will not survive the extreme heating environment nestled between two SRB's. The baseline RS-68 is ablatively cooled, and apparently only a regeneratively-cooled engine can tough it out. The DIRECT team settled for throwing away SSME's as their core engine, believing that this was more cost-effective than developing the proposed "regen" RS-68R. But as I've argued in my previous post, SSME's that are designed for low production costs will be very different from the baseline SSME.
A regen nozzle for RS-68 is simple in concept. Keep the inner mold line the same (to preserve the expansion ratio,) and manufacture the nozzle from thinner stock (since the extra thickness won't be needed for ablating away during ascent.) Mill some cooling channels into the outer nozzle, then braze a thin cooling jacket over the top. Ironically, I'd suspect that the baseline RS-68 shows slight improvement in Isp the longer it burns, because the expansion ratio increases as more material is burned off the inner surface of the nozzle. RS-68R will not experience the same effect.
The turbopumps need to be analyzed to ensure they can pump enough cryogenically-chilled propellant through the coolant channels. RS-68 already employs regen cooling in the combustion chamber, so adding it to the nozzle shouldn't be a huge challenge. Still, this may require a further upgrade beyond what's planned for RS-68A's turbopumps. Previous studies of RS-68R, with the same expansion ratio as the current RS-68, show an Isp improvement from 409 sec. to 419 sec.
RS-68R is definitely a step in the right direction for improving the engine's mass, durability and performance. Still, it falls short of the SSME's 453 second Isp. A longer, wider nozzle would improve specific impulse, at the expense of lower thrust. Regardless of whether RS-68R's nozzle ever reached the expansion ratio of the SSME, I'd never expect to duplicate SSME's specific impulse. After all, SSME makes use of the more complex staged combustion cycle, and has a much higher chamber pressure. Still, it's not unreasonable to expect an Isp around 430 seconds for RS-68R if the nozzle is redesigned to a higher expension ratio. This is an estimate for how the proposed Space Transportation Main Engine would have performed; it would have used an expansion ratio somewhere between RS-68 and SSME, had a chamber pressure between those two extremes, and used RS-68's gas generator cycle.
At the end of the day, NASA-Marshall will have some big decisions to make, and the trades should be backed up by something more substantial and detailed than a 60-day study. Will it be more cost-effective to develop expendible SSME or RS-68R? And if RS-68R is cheaper to develop, procure on a unit basis, or both, will those cost-savings be offset by the engine's lower efficiency when comapred to expendible SSME? The last metric can be quantified with the additional number of launches that will be required to put the same payload mass in space.
The current evolutionary roadmap for RS-68 goes to RS-68A, funded by DoD to improve performance for the Delta IV Heavy and eventually the entire Delta IV fleet. By upgrading the turbopumps for higher flowrates, PWR will put more injectors into the thrust chamber of RS-68A to improve thrust and specific impluse (raising from 409 sec to 414 sec in vacuum.)
NASA's current roadmap relied on a further upgrade of RS-68, the RS-68B model, for the Ares V super-booster. The NASA-funded upgrade would increase the engine's redundancy in certain systems (required by the current man-rating rules, and a focal point of the Delta IV crew-launch studies) and reducing the amount of gas that collects at the launch pad prior to ignition.
Yet the findings of the DIRECT team prior to unveiling DIRECT 3.0 may put a wrench in NASA's plans. Their studies (and apparently NASA's internal studies confirm this) indicate the current RS-68 engines will not survive the extreme heating environment nestled between two SRB's. The baseline RS-68 is ablatively cooled, and apparently only a regeneratively-cooled engine can tough it out. The DIRECT team settled for throwing away SSME's as their core engine, believing that this was more cost-effective than developing the proposed "regen" RS-68R. But as I've argued in my previous post, SSME's that are designed for low production costs will be very different from the baseline SSME.
A regen nozzle for RS-68 is simple in concept. Keep the inner mold line the same (to preserve the expansion ratio,) and manufacture the nozzle from thinner stock (since the extra thickness won't be needed for ablating away during ascent.) Mill some cooling channels into the outer nozzle, then braze a thin cooling jacket over the top. Ironically, I'd suspect that the baseline RS-68 shows slight improvement in Isp the longer it burns, because the expansion ratio increases as more material is burned off the inner surface of the nozzle. RS-68R will not experience the same effect.
The turbopumps need to be analyzed to ensure they can pump enough cryogenically-chilled propellant through the coolant channels. RS-68 already employs regen cooling in the combustion chamber, so adding it to the nozzle shouldn't be a huge challenge. Still, this may require a further upgrade beyond what's planned for RS-68A's turbopumps. Previous studies of RS-68R, with the same expansion ratio as the current RS-68, show an Isp improvement from 409 sec. to 419 sec.
RS-68R is definitely a step in the right direction for improving the engine's mass, durability and performance. Still, it falls short of the SSME's 453 second Isp. A longer, wider nozzle would improve specific impulse, at the expense of lower thrust. Regardless of whether RS-68R's nozzle ever reached the expansion ratio of the SSME, I'd never expect to duplicate SSME's specific impulse. After all, SSME makes use of the more complex staged combustion cycle, and has a much higher chamber pressure. Still, it's not unreasonable to expect an Isp around 430 seconds for RS-68R if the nozzle is redesigned to a higher expension ratio. This is an estimate for how the proposed Space Transportation Main Engine would have performed; it would have used an expansion ratio somewhere between RS-68 and SSME, had a chamber pressure between those two extremes, and used RS-68's gas generator cycle.
At the end of the day, NASA-Marshall will have some big decisions to make, and the trades should be backed up by something more substantial and detailed than a 60-day study. Will it be more cost-effective to develop expendible SSME or RS-68R? And if RS-68R is cheaper to develop, procure on a unit basis, or both, will those cost-savings be offset by the engine's lower efficiency when comapred to expendible SSME? The last metric can be quantified with the additional number of launches that will be required to put the same payload mass in space.
posted by Mr. X at
3:50 PM
