Quote from: aero on 12/16/2013 10:48 pmDoes anyone know what the actual payload mass was for the Cassiope launch? what was it?If not, I've searched and found the satellite masses but not the mass of any ancillary parts.Cassiope - 481 kgCU Sat - 40.82 kgDANDE - 50 kgDOPACS - 4.5 kgDoes anyone know what the mass of the payload adapter? What was it?I know the trajectory of the first stage up to MECO pretty well from repeatedly viewing the launch (available on livestream), pausing and recording the data as it is given. Using this stage 1 trajectory I conclude that the F9.1 could not reach LEO with 13,150 kg though it can with other trajectories. I'm trying to simulate the boost-back so knowing the payload would help determine the fuel available for that function.Cassiope went to a 300 x 1,500 km x 80 deg orbit from Vandenberg AFB. The 13.15 tonne figure is given for a 185 km x 28.5 deg orbit from Cape Canaveral, where extra delta-v is provided by the Earth's rotation. Payload from Vandenberg toward a near-polar orbit will be much less - and the higher altitude will take away even more lifting capability. Falcon 9 v1.1 is probably only able to lift 9 tonnes or less to the Cassiope orbit, assuming that the advertised capabilities are accurate. - Ed Kyle
Does anyone know what the actual payload mass was for the Cassiope launch? what was it?If not, I've searched and found the satellite masses but not the mass of any ancillary parts.Cassiope - 481 kgCU Sat - 40.82 kgDANDE - 50 kgDOPACS - 4.5 kgDoes anyone know what the mass of the payload adapter? What was it?I know the trajectory of the first stage up to MECO pretty well from repeatedly viewing the launch (available on livestream), pausing and recording the data as it is given. Using this stage 1 trajectory I conclude that the F9.1 could not reach LEO with 13,150 kg though it can with other trajectories. I'm trying to simulate the boost-back so knowing the payload would help determine the fuel available for that function.
Don't you also incur a penalty for Cassiope's use of a single upper stage burn vs. double upper stage burn trajectory.Though to be fair, if the payload needs the extra boost, there is no reason they would not use a two burn profile.
The problem with having the Falcon second stage do three burns (LEO injection, raise apogee to GEO, raise perigee to GEO also) is the 5-6 hours spent between the second and third orbit. Some of the engine lines froze in the CASSIOPE mission in less than an hour. The SES launch handled the half-hour wait fine, but 5-6 hours is a lot longer. On top of that is the issue of keeping the LOX tank stable during that time. If I'm not mistaken, those considerations were why they invented the current technique, the one used in the SES launch.
I know they didn't invent this technique for this mission, and that the technique has been around for a very long time. But it wasn't there in the beginning (the 60s), and that was where I was coming from. I wasn't fully aware of the current technique until I saw it happening with the SES launch. It flew beneath my personal radar in the years since Apollo. I wasn't paying attention to the techniques of GEO satellite launching, being distracted by moon launches, space shuttles and interplanetary probes...
The "techniques of GEO satellite launching" were developed from "moon launches, space shuttles and interplanetary probes" in the 60's. The two burns were done for Surveyor in 1967 by Atlas Centaur and for Apollo 8 in 1968.
Quote from: Jim on 12/17/2013 02:32 pmThe "techniques of GEO satellite launching" were developed from "moon launches, space shuttles and interplanetary probes" in the 60's. The two burns were done for Surveyor in 1967 by Atlas Centaur and for Apollo 8 in 1968.That's not the same thing, and you know it. Both of those launches were translunar injection, which is not the same thing as coming up with a streamlined, efficient technique for launching a satellite to geosynchronous orbit.
I don't know that anybody in the 60s would have ever dreamed of what they have been doing around Saturn these past ten years with the Cassini probe, or what they did in the 90s with the Galileo probe around Jupiter.
Yes, they did. The sling shot maneuver was proposed in 1961 and used by Mariner 10 in 1973 (which means it was designed in the 60's).
I don't know when somebody got the bright idea of having the final kick motor on the satellite itself, as a sort of final stage, using hypergolics or something else that keeps well. The satellite needed it anyway, for station-keeping. But I strongly suspect that they didn't think right away of starting off with an apogee 50% higher than GEO so that the essential energy level of the orbit starts off more or less the same as the target orbit.
Quote from: Jim on 12/17/2013 06:12 pmYes, they did. The sling shot maneuver was proposed in 1961 and used by Mariner 10 in 1973 (which means it was designed in the 60's).Slingshot after slingshot after slingshot, for years on end?
While I will agree that there is no essential difference between the second burn of Falcon and that of the other rockets you referred to, what I was originally referring to was the notion (long ago) that the final insertion into GTO would be handled by that last stage, making a third burn in total. That is the burn I was referring to as presenting the freezing or pressurization difficulties.
Since day 1 (going back to Syncom 1, the first GSO satellite). All the Delta and Atlas spacecraft going to GSO had the kick motor in them. All commercial comsats have kick motors or boost systems. Only the DOD spacecraft on Titan IIIC, 34D and IV using Transstage, IUS, and Centaur and Delta IV Heavy have relied on the launch vehicle for final injection into GSO, which is the exception vs the rule.
Which begs the question: given that using a kick motor that is part of the satellite is more efficient and practical, why would the DOD use the final stage that way at all? Not arguing here. Just curious and learning stuff...