On Aug 7, 4:28*am, Robert Clark wrote:

*The Russian RD-171 and RD-180 are high thrust, liquid-fueled engines
still in active operation. Here's the Astronautix pages on the RD-171
and RD-180:

RD-171http://www.astronautix.com/engines/rd171.htm

RD-180http://www.astronautix.com/engines/rd180.htm

*You would need 2 to 3 of these to match the thrust of the Ares I
first stage solids. But the thrust to weight is so much better you
might be able to match the payload to orbit just using one of these
engines. (You would have 200,000 lbs less dry mass at launch.)



This page gives altitude and velocity of the Ares I at first stage
separation as 59 km and 2024 m/s:

Space Launch Report - Ares I.
http://www.spacelaunchreport.com/ares1.html

Then we can estimate how much fuel it would take to reach this delta-
v, including the gravity drag for that altitude, based on the Isp of
the liquid fuel engines and the mass of the Ares upper stage.

A preliminary calculation shows the RD-180 wouldn't have enough
thrust for the fuel load required. However, the RD-171 should be able
to do it using the specifications given he

RD-171
http://www.astronautix.com/engines/rd171.htm

The second stage of the Ares I is about 175,000 kg, when you include
payload. The RD-171 weighs 9,500 kg. Even if you used as much fuel
mass as the SRB of 1,400,000 lb, a tankage mass ratio of 1/100th the
propellant mass for kerosene/LOX engines would only add 7,000 kg. So
call the the upper stage plus the empty weight of the lower stage
200,000 kg.
You want to reach the same velocity of 2,000 m/s and altitude of 50
km reached by the SRB. The delta-v required for the altitude can be
found from the equation v^2 = 2gh. So for h = 50,000 m, v = 990 m/s.
Air drag losses it turns out are relatively small for large
cylindrical rockets that get rapidly out of the atmosphere, about 150
m/s for medium sized launchers by this page:

Flight Mechanics of Manned Sub-Orbital Reusable Launch Vehicles with
Recommendations for Launch and Recovery.
http://www.spacefuture.com/archive/f... ecovery.shtml

So let's call the total delta-v 3,000 m/s. The Isp of the RD-171 is
given on the Astronautix page as 309 s at sea level and 337 s in
vacuum. Let's give it an average Isp of 320 s. Then the mass ratio is
exp(3,000/3200) = 2.554. So for a mass at first stage burnout of
200,000 kg, the mass at launch would be 2.554x200,000 = 510718 kg. Of
this 510,718-200,000 kg = 310,718 kg would be fuel, less than half of
the fuel load of the SRB for Ares I.
The launch mass of 510,718 kg = 1,123,580 lb is well within the
thrust capabilities of the RD-171 to lift, with its sea level thrust
of 1,697,300 lb.


Bob Clark