SpaceX Falcon 9 - Inmarsat 5 F4 - May 15, 2017 - DISCUSSION

[Dante80]

Perhaps now I actually understand why people are so reluctant to even entertain the possibility that F9 expendable with zero margins might be able to put an 8.3 ton payload in GTO-1800 m/s. That would challenge every rocket in service, except for Ariane V and D4H.
It really must be hard to conceive that such a cheap rocket can get that much performance. But I'm a believer, eventually there will be one Block V expendable launch that will place something like a 7.5 ton payload to an orbit similar to this, and awe the world ! A payload large enough that in requires the same effort to put that 8.3 tons to GTO-1800 !

Actually, can anyone calculate what Ariane 5's payload would be if it were launched from 28 degrees?

Wouldn't it be the same (10,500kg) but to GTO-1800 instead of GTO-1500?

[wannamoonbase]

Ultimately, it took a year and a half, including a freak pad accident to surmount this challenge. Judging from the last two campaigns though, I think it was worth it in the end.

Yeah it does, the last two rockets look like they were supercharged.

Still need a lot more launches before getting cocky.  But this it appears to be a big edge, looking forward to the Block 4 and 5.

[woods170]

The machine has evolved, and the most recent two flights have exhibited a new level of performance - to the extent that I'm convinced we are seeing at least a Block 4 second stage.

Not sure about two latest flights, but the last one was totally out of family. Looks like a different rocket.

It is. No recovery hardware on stage 1 and different mission profile for stage 1.

Supercharged? No. Not yet.

[hans_ober]

Why don't they always do minimal residual shutdowns?

They can launch a 5t sat to GTO-1770 with a normal shutdown, but why not let it burn a little longer and use all the fuel?

Let's say that they're being conservative (and have a fuel margin in case of underperformance) and quote that they can launch the 5t sat to GTO-1800.
During the launch, all goes well and this reserve propellant isn't actually needed, so why not use it and burn to a higher apogee, which takes the satellite closer.

Also, what is the actual difference in dv for plane change vs higher apogee? Any graph for reference?

[vanoord]

Why don't they always do minimal residual shutdowns?

They can launch a 5t sat to GTO-1770 with a normal shutdown, but why not let it burn a little longer and use all the fuel?

Let's say that they're being conservative (and have a fuel margin in case of underperformance) and quote that they can launch the 5t sat to GTO-1800.
During the launch, all goes well and this reserve propellant isn't actually needed, so why not use it and burn to a higher apogee, which takes the satellite closer.

Also, what is the actual difference in dv for plane change vs higher apogee? Any graph for reference?

There's an argument that saving a bit of fuel to de-orbit the stage mightn't be a bad idea.

[rpapo]

There's an argument that saving a bit of fuel to de-orbit the stage mightn't be a bad idea.

And now that the upper stage seems to have a longer life, why not wait until apogee to do the deorbit burn?  From 60,000km, very little fuel would be needed to lower the 400km perigee down into the atmosphere.

[rsdavis9]

May I ask how high an apogee would this launch have produced if it did zero inclination reduction ?
Perhaps this will better explain to people how significant this performance was, considering it was also the heaviest F9 GTO payload to date !

The problem with orbital mechanics is that mixing apogee increase AND plane change is a lot cheaper than doing one and then the other one. I will yield to Lou to do such calculation, though.

I thought that the best usage of LV performance was to put the GTO payload on as high as possible apogee, and THEN once its on a super sync trajectory, it can use the lower speeds of the apogee to effect some inclination change and reduction in apogee on each orbital apogee and increase in perigee on each orbital perigee, done by the payload itself.

That led me to think that if SpaceX could create a mini ITS rocket that had perhaps 5+ days of mission endurance, it could do a bi elliptical transfer by itself, by going into a super sync orbit, doing the entire inclination change and apogee reduction in a single burn, then the perigee raising in the other half orbit and deliver a large number of GEO payloads into GEO-500m/s with zero inclination and just some circularization left, so the orbital period is a few hours away from GEO, so the payloads can pace themselves to go directly into their exact slots, although they would all be delivered to the same initial orbit.
The mini ITS would then do the required orbital transfer to re-enter and land, avoiding brute force trajectory corrections to get to the LZ.

If you didn't have the extra mass of S2 it is not overall the most efficient to go to a super synchronous(SS). SS works because S2 expends more so the lighter satellite has less work to do. If there were no stages or no engines on the satellite the most efficient is:
1. Launch to parking
2. boost apogee at the equator to GEO.
3. boost perigee and remove inclination at apogee.
 
I have been playing with the equations.
   https://en.wikipedia.org/wiki/Geostationary_transfer_orbit
   https://en.wikipedia.org/wiki/Hohmann_transfer_orbit
I have made a bc(unix) program for doing the calcs. I have yet to incorporate the excellent calc from Lou.
   https://forum.nasaspaceflight.com/index.php?topic=36954.0

[rsdavis9]

There's an argument that saving a bit of fuel to de-orbit the stage mightn't be a bad idea.

And now that the upper stage seems to have a longer life, why not wait until apogee to do the deorbit burn?  From 60,000km, very little fuel would be needed to lower the 400km perigee down into the atmosphere.

So what delta V can the GN2 thrusters do?
Maybe it is enough?

EDIT:

I get 20 m/s for the perigee lowering from 380km to 100km.
orbvel of 380x60000 = 1053 m/s
orbvel of 100x60000=1033 m/s

[Star One]

...

Zenit-3SL can do 6.16 to a 1,477m/s deficit GTO. That's a ~95m/s difference. If they used less delta-v GTO, and they didn't had structural limits on the rocket, it would be much higher performance. Using a linear approximation I get 7.8 tonnes.

Isn't Zenit retired or on last launch?

I am not so sure on that. I certainly wouldn't recommend you go to Ukraine and say that.

[rpapo]

So what delta V can the GN2 thrusters do?
Maybe it is enough?

You were thinking along the same lines, but I somehow find it hard to imagine the GN2 system having enough propellant and energy to accelerate the second stage (4500kg?) by 20 m/s.

[gospacex]

...

Zenit-3SL can do 6.16 to a 1,477m/s deficit GTO. That's a ~95m/s difference. If they used less delta-v GTO, and they didn't had structural limits on the rocket, it would be much higher performance. Using a linear approximation I get 7.8 tonnes.

Isn't Zenit retired or on last launch?

I am not so sure on that. I certainly wouldn't recommend you go to Ukraine and say that.

Why not? It's not like people in Yuzhmash are oblivious to the fact that RD-170 is not made in Ukraine.

[Star One]

...

Zenit-3SL can do 6.16 to a 1,477m/s deficit GTO. That's a ~95m/s difference. If they used less delta-v GTO, and they didn't had structural limits on the rocket, it would be much higher performance. Using a linear approximation I get 7.8 tonnes.

Isn't Zenit retired or on last launch?

I am not so sure on that. I certainly wouldn't recommend you go to Ukraine and say that.

Why not? It's not like people in Yuzhmash are oblivious to the fact that RD-170 is not made in Ukraine.

I thought they were doing their upmost to keep Zenit going.

[rsdavis9]

So what delta V can the GN2 thrusters do?
Maybe it is enough?

You were thinking along the same lines, but I somehow find it hard to imagine the GN2 system having enough propellant and energy to accelerate the second stage (4500kg?) by 20 m/s.

So to continue with my calcs
if the isp is 200s for GN2?
we get a mass fraction of 1.01 or 1/100 of the 4500kg which is 45 kg.
if the isp is 300 it is 1.007 or .006*4500=27kg

EDIT:
Unknowns are
isp of GN2 thrusters?
kg of GN2 on board S2?

[gospacex]

...

Zenit-3SL can do 6.16 to a 1,477m/s deficit GTO. That's a ~95m/s difference. If they used less delta-v GTO, and they didn't had structural limits on the rocket, it would be much higher performance. Using a linear approximation I get 7.8 tonnes.

Isn't Zenit retired or on last launch?

I am not so sure on that. I certainly wouldn't recommend you go to Ukraine and say that.

Why not? It's not like people in Yuzhmash are oblivious to the fact that RD-170 is not made in Ukraine.

I thought they were doing their upmost to keep Zenit going.

How? Ukraine will not use a Russian engine. It's a political suicide to anyone to even propose that.

[Celestar]

The theoretical limit of cold gas (N2) thrusters seems to be around 80s [1]

[1] http://cdn.intechopen.com/pdfs-wm/37528.pdf

Celestar

[Kaputnik]

So what delta V can the GN2 thrusters do?
Maybe it is enough?

You were thinking along the same lines, but I somehow find it hard to imagine the GN2 system having enough propellant and energy to accelerate the second stage (4500kg?) by 20 m/s.

So to continue with my calcs
if the isp is 200s for GN2?
we get a mass fraction of 1.01 or 1/100 of the 4500kg which is 45 kg.
if the isp is 300 it is 1.007 or .006*4500=27kg

EDIT:
Unknowns are
isp of GN2 thrusters?
kg of GN2 on board S2?

Way off. isp of GN2 is more like 70s.

https://en.m.wikipedia.org/wiki/Cold_gas_thruster

[rsdavis9]

So what delta V can the GN2 thrusters do?
Maybe it is enough?

You were thinking along the same lines, but I somehow find it hard to imagine the GN2 system having enough propellant and energy to accelerate the second stage (4500kg?) by 20 m/s.

So to continue with my calcs
if the isp is 200s for GN2?
we get a mass fraction of 1.01 or 1/100 of the 4500kg which is 45 kg.
if the isp is 300 it is 1.007 or .006*4500=27kg

EDIT:
Unknowns are
isp of GN2 thrusters?
kg of GN2 on board S2?

Way off. isp of GN2 is more like 70s.

https://en.m.wikipedia.org/wiki/Cold_gas_thruster

so isp of 70s
1.029
.029*4500=130kg
Sounds like a lot.

[LouScheffer]

Why don't they always do minimal residual shutdowns?

There are some practical considerations:
(a) The satellite has to be able to work when above GEO.  Not all satellites are designed to do this.
(b) With a targeted shutdown, maneuvers can be planned and verified in advance, and you know what ground stations are needed and where they should point.  (e.g, on Tuesday, at 4PM, we'll raise the perigee, using station X for commanding)  With minimum residual, maneuvers must be computed in real time - you need to see what the rocket does.  Plus they are more complex, at least two burns.
(c) Trajectory calculations are more complex for super-synchronous.   The moon's influence may need to be included.  One early super-sync mission (SuperBird 6) failed since the moon perturbed the orbit enough to give a too-low perigee for the transfer orbit.

Also, what is the actual difference in dv for plane change vs higher apogee? Any graph for reference?

Plot from this post.

[wannamoonbase]

So what delta V can the GN2 thrusters do?
Maybe it is enough?

You were thinking along the same lines, but I somehow find it hard to imagine the GN2 system having enough propellant and energy to accelerate the second stage (4500kg?) by 20 m/s.

20 m/s would be how long a burn for a MVac at minimum throttle, 0.5 seconds?

[rpapo]

So what delta V can the GN2 thrusters do?
Maybe it is enough?

You were thinking along the same lines, but I somehow find it hard to imagine the GN2 system having enough propellant and energy to accelerate the second stage (4500kg?) by 20 m/s.

20 m/s would be how long a burn for a MVac at minimum throttle, 0.5 seconds?

Quite likely.  The real problem with that would simply be one of how to pull off such a short, precise burn, especially when you count in the preliminary steps, like ullage thrusting and so on.  The Merlin Vacuum is almost too powerful a tool for the job, while the nitrogen thrusters don't seem to be enough.

There would also be targeting accuracy to take into account here.  A small difference in a retro-burn at 60,000 km could make a large difference in where reentry will take place.  The deep south Pacific or Indian Ocean are big places, but even so...