References
Encyclopedia Astronautica by Mark Wade

F-1

Ready: 1966
Burn Time: 161 seconds
Propellants: LOX/RP-1 Kerosene
Thrust (sl): 1,516,000 lbf
Thrust (vac): 1,740,000 lbf
ISP (sl): 265
ISP (vac): 304
Weight: 18,500 lbs
Oxidizer to Fuel Ratio: 2.27
Fuel Density: 1,012 kg/m3
T/W Ratio (sl): 81.95 lbf per lb
T/W Ratio (vac): 94.05 lbf per lb

F-1A

Ready: 1970
Burn Time: 158 seconds
Propellants: LOX/RP-1 Kerosene
Thrust (sl): 1,800,000 lbf
Thrust (vac): 2,000,000 lbf
ISP (sl): 270
ISP (vac): 310
Weight: 17,853 lbs
T/W Ratio (sl): 100.82 lbf per lb
T/W Ratio (vac): 112.03 lbf per lb

J-2

Ready: 1966
Burn Time: 475 seconds
Propellants: LOX/LH2
Thrust (sl): 109,300 lbf
Thrust (vac): 232,250 lbf
ISP (sl): 200
ISP (vac): 421
Weight: 3,170 lbs
Oxidizer to Fuel Ratio: 5.5
Fuel Density: 344 kg/m3
T/W Ratio (sl): 34.48
T/W Ratio (vac): 73.26

Aerojet-General M-1

Ready: 1970s
Propellants: LOX/LH2
Thrust (sl): 868,851 lbf
Thrust (vac): 1,199,575 lbf
ISP (sl): 310
ISP (vac): 428
Weight: 19,991.5 lbs
Oxidizer to Fuel Ratio: 6
Fuel Density: 362 kg/m3
T/W Ratio (sl): 43.46
T/W Ratio (vac): 60

Notes: Was cancelled in 1965, due to the general draw down of NASA and the elimination of post-Apollo missions and hardware development.

Rocketdyne RS-25; aka Space Shuttle Main Engine (SSME)

Ready: 1980~
Burn Time: 480 seconds
Propellants: LOX/LH2
Thrust (sl): 408,570 lbf
Thrust (vac): 512,100 lbf
ISP (sl): 363
ISP (vac): 453
Weight: 7,000 lbs
Oxidizer to Fuel Ratio: 6
Fuel Density: 362 kg/m3
T/W Ratio (sl): 58.37 lbf per lb
T/W Ratio (vac): 73.16 lbf per lb

Rocketdyne RS-84

Ready: 2000~
Propellants: LOX/RP-1
Thrust (sl): 1,064,000 lbf
Thrust (vac): 1,130,000 lbf
ISP (sl): 304
ISP (vac): 324
Weight: 17,919 lbs
Oxidizer to Fuel Ratio: 2.7
Fuel Density: 1,025.18 kg/m3
T/W Ratio (sl): 59.38 lbf per lb
T/W Ratio (vac): 63.06 lbf per lb

Notes: Developed as a reusable LOX/RP-1 engine with a lifetime of about 100 flights. Cancelled by NASA in 2005.

Rocketdyne RS-68

Ready: 2000~
Propellants: LOX/LH2
Thrust (sl): 663,000 lbf
Thrust (vac): 758,000 lbf
ISP (sl): 359
ISP (vac): 409
Weight: 14,876 lbs
Oxidizer to Fuel Ratio: 6
Fuel Density: 362 kg/m3
T/W Ratio (sl): 44.57 lbf per lb
T/W Ratio (vac): 50.95 lbf per lb

Notes: Developed as a cheap expendable engine building off the failed STME program. Has 80% fewer parts and requires 92% less labor than than the RS-25 SSME. It uses an ablative engine nozzle, which while being heavier than a regeneratively cooled nozzle, is much cheaper and faster to construct than a regenerative nozzle with it's hordes of tubing.

J-2X

Ready: 2012~
Burn Time: 431 seconds
Propellants: LOX/LH2
Thrust (vac): 294,490 lbf
ISP (vac): 448
Weight: 5,450 lbs
Oxidizer to Fuel Ratio: 5.5
Fuel Density: 344 kg/m3
T/W Ratio (vac): 54.03 lbf per lb

Note: Had to meet much more stringent requirements, including a restart in space after 90~ days cold.

NERVA Model 2

Ready: 1970s~
Burn Time: 1,200 Seconds
Propellants: Nuclear/LH2
Thrust (sl): 89,810 lbf
Thrust (vac): 195,000 lbf
ISP (sl): 380
ISP (vac): 825
Weight: 26,140 lbs
Fuel Density: 71 kg/m3
T/W Ratio (sl): 3.44 lbf per lb
T/W Ratio (vac): 7.46 lbf per lb

Timberwind 250

Ready: 1990s~
Burn Time: 493 Seconds
Propellants: Nuclear/LH2
Thrust (sl): 429,900 lbf
Thrust (vac): 551,140 lbf
ISP (sl): 780
ISP (vac): 1,000
Weight: 18,200 lbs
Fuel Density: 71 kg/m3
T/W Ratio (sl): 23.62 lbf per lb
T/W Ratio (vac): 30.28 lbf per lb

Notes: Designed as a pebble-bed type reactor by DARPA/DoD for advanced nuclear payloads from Shuttle in support of SDI. Development ended in 1992.

Aerojet BPT-4000 Hall Effect Thruster

Ready: 2000s
Burn Time: 6,000+ Hours
Propellants: Electric/Xenon
Thrust (vac): 0.06 lbf (0.27 newtons)
ISP (vac): 1,950
Weight: 16.5 lbs
Fuel Density: 3,100 kg/m3 plus 4500 W at 350 V
T/W Ratio (vac): 0.0036 lbf per lb

NASA Evolutionary Xenon Thruster (NEXT)

Ready: 2010s
Propellants: Electric/Xenon
Thrust (vac): 0.05 lbf (0.24 newtons)
ISP (vac): 4,100~
Fuel Density: 3,100 kg/m3 plus unknown amount of electric power.

NASA High Power Electric Propulsion (HiPEP)

Ready: 2010s
Propellants: Electric/Xenon
Thrust (vac): 0.15 lbf (0.67 newtons)
ISP (vac): 9,620~
Fuel Density: 3,100 kg/m3 plus unknown amount of electric power.

Ad Astra VX-200 VASMIR Engine

Ready: 2015-2020?
Propellants: Electric/Argon
Thrust (vac): 0.89 lbf (4 newtons)
ISP (vac): 6,000
Weight: about 661 lbs (300 kg) (based on 1.5 kg/kW specific power)
Fuel Density: 1,430 kg/m3 plus 200~ kW of electric power
T/W Ratio (vac): about 0.0013 lbf per lb

Ad Astra Theoretical 10 Megawatt VASMIR Engine for Interplanetary Flight

Ready: 2020-2030?
Propellants: Electric/Argon
Thrust (vac): 43~ lbf (191~ newtons)
ISP (vac): 6,000
Weight: about 33,069 lbs (15,000 kg) (based on 1.5 kg/kW specific power)
Fuel Density: 1,430 kg/m3 plus 10 MW of electric power
T/W Ratio (vac): about 0.0013 lbf per lb

Notes: This has been proposed by Ad Astra; and the specifications are based off of the specific power and weights of their existing VX-200. No doubt during the time of development between the VF-200 flight engine set to fly on the ISS in 2010-2011, and when we reach the point of deploying a 10 megawatt VASMIR engine, specific weights and power will improve.