SpaceX Falcon 9 - EchoStar 23 - March 16, 2017 - DISCUSSION

[Steven Pietrobon]

(I had asked this, then found it later ) Please provide a link for the technical webcast that provided those inside-the-tank views.

Here is the technical webcast where I got the captures from.

[Jarnis]

Here are some screen captures of the roll out.

Not Echostar 23 rollout. Legs. Dragon.

Still interesting and proves it really is a rocket that has its nose pointing to the ground during nominal operation. Usually when the nose points at the ground, things have gone REALLY bad. 

[jpo234]

In the Technical Webcast were some frames from the second stage LOX camera. They show a white (??) pressure vessel that is absent from previous tank camera images. Any idea what this is?
Experimental COPV replacement?

[Kaputnik]

Just realised that with this launch, F9 becomes the most launched vehicle of the year- although Atlas will match it again when OA7 launches.

[stcks]

In the Technical Webcast were some frames from the second stage LOX camera. They show a white (??) pressure vessel that is absent from previous tank camera images. Any idea what this is?
Experimental COPV replacement?

This is nothing new. We've seen this before a few times, notably on SES-9: http://spaceflight101.com/falcon-9-ses-9/wp-content/uploads/sites/71/2016/03/SES9-11.jpg

[cscott]

In the Technical Webcast were some frames from the second stage LOX camera. They show a white (??) pressure vessel that is absent from previous tank camera images. Any idea what this is?
Experimental COPV replacement?

This is nothing new. We've seen this before a few times, notably on SES-9: http://spaceflight101.com/falcon-9-ses-9/wp-content/uploads/sites/71/2016/03/SES9-11.jpg

Maybe a GSO kit? Allows extended coast, something like that?

[LouScheffer]

On this flight, compared to SES-9, the first stage seemed to have similar performance, but the second stage seemed upgraded.

Compared to SES-9, this first stage burned 6-7 seconds longer.  This makes sense since SES-9 had about 17 second,  3 engine, entry burn and 5x3 landing burn.  Also, the staging speed is 368 m/s greater (9624 km/hr vs 8300).  This is about 7 seconds at 5 Gs, so also compatible.

The second stage had a significantly shorter burn time, however.   Compared to SES-9, it burned about 24 seconds less in total.  It started 6 seconds later, but finished its first burn 22 seconds earlier (8:33 as opposed to 8:55).  The GTO injection burn was about 4 seconds longer.  Assuming they had the same fuel load, it chomped through it about 24/420, or 6%, faster.   This would equate to 6% higher thrust.

[LouScheffer]

Anyone know the GTO transfer orbit?  I could not find it, at least with a simple search.

I'd assume they would use the extra performance from expending the first stage to help the payload.

It looks like the first stage provided 360 m/s more.  If they use 300 m/s of this for the injection burn (they are not burning to depletion), they could:

(a) go super-sync to an apogee of 67,000 km, reducing deficit to 1630 m/s, or
(b) reduce inclination to about 20 degrees, with GEO apogee, reducing deficit to 1660 m/s or so.

I'd guess they did (b), since that does not need any additional trajectory work by the customer.

[IntoTheVoid]

On this flight, compared to SES-9, the first stage seemed to have similar performance, but the second stage seemed upgraded.

Compared to SES-9, this first stage burned 6-7 seconds longer.  This makes sense since SES-9 had about 17 second,  3 engine, entry burn and 5x3 landing burn.  Also, the staging speed is 368 m/s greater (9624 km/hr vs 8300).  This is about 7 seconds at 5 Gs, so also compatible.

The second stage had a significantly shorter burn time, however.   Compared to SES-9, it burned about 24 seconds less in total.  It started 6 seconds later, but finished its first burn 22 seconds earlier (8:33 as opposed to 8:55).  The GTO injection burn was about 4 seconds longer.  Assuming they had the same fuel load, it chomped through it about 24/420, or 6%, faster.   This would equate to 6% higher thrust.

Until we know the orbit achieved, doesn't it make more sense to presume that the 2nd stage burned shorter not due to upgraded performance, but because the 1st stage burned longer and therefore a shorter 2nd stage burn was needed for the contracted orbit?

[abaddon]

I'd assume they would use the extra performance from expending the first stage to help the payload.

IIRC Jim has said in the past that this is not necessarily the case.  It depends on the payload and the trajectory planning.  I am pretty sure we have seen both cases before.  From what you are saying (shorter burn than SES-9) my guess would be that it (S2) hit the target parameters and then shut down, rather than burning to depletion.  That's a WAG though.

[Jim]

If so, why did they do that rather than stay parallel?

Spacecraft requirements

[LouScheffer]

I'd assume they would use the extra performance from expending the first stage to help the payload.

IIRC Jim has said in the past that this is not necessarily the case.  It depends on the payload and the trajectory planning.  I am pretty sure we have seen both cases before.  From what you are saying (shorter burn than SES-9) my guess would be that it (S2) hit the target parameters and then shut down, rather than burning to depletion.  That's a WAG though.

This is certainly possible - they could have simply stopped with 25 seconds of fuel left, when they reached their target orbit.  It's a lot of performance to leave on the table, though.  They could have saved at least 100 m/s of delta-V needed to reach GEO, which roughly equates to 2 years of satellite station-keeping lifetime. 

One way I could see this happening is is SpaceX thought they could do recovery.   In this case they could only do a 28 degree, GTO-1800 m/s transfer orbit.  Then when they switched to expendable, the customer had already made plans for this orbit and said improving the orbit was not worth redoing the calculations and renegotiating the contract.  So they left the old orbit and just had excess performance.

We'll know more once we get the transfer orbit specs.

[kevin-rf]

Anyone know the GTO transfer orbit?  I could not find it, at least with a simple search.

No TLE's on Space-Trak yet.

Latest TLE is 42069U which appears to have come from 98067LE, aka ISS (Norad 25544U) . That or I'm reading the TLE's wrong.

[HVM]

In the Technical Webcast were some frames from the second stage LOX camera. They show a white (??) pressure vessel that is absent from previous tank camera images. Any idea what this is?
Experimental COPV replacement?

Due lighting and sharpness' I would guess that, it's something small and close to the camera than big/far.

[wannamoonbase]

Great flight, now bring on the next one!!

[mn]

Noticed this interesting text in the technical webcast.

I know it's the technical webcast, but I don't know if they need to get that technical

[abaddon]

Yeah, I watched the broadcast this morning, and saw that as well.  Oops .

I guess if you're going to get something wrong, it is a good way to go.

[sunbingfa]

Anyone know the GTO transfer orbit?  I could not find it, at least with a simple search.

I'd assume they would use the extra performance from expending the first stage to help the payload.

It looks like the first stage provided 360 m/s more.  If they use 300 m/s of this for the injection burn (they are not burning to depletion), they could:

(a) go super-sync to an apogee of 67,000 km, reducing deficit to 1630 m/s, or
(b) reduce inclination to about 20 degrees, with GEO apogee, reducing deficit to 1660 m/s or so.

I'd guess they did (b), since that does not need any additional trajectory work by the customer.

It seems you got it right, man! I've got to learn more.......

42070   ECHOSTAR 23   2017-014A     632.93min   22.43deg   35903km X   179km
 

[Space Ghost 1962]

Anyone know the GTO transfer orbit?  I could not find it, at least with a simple search.

I'd assume they would use the extra performance from expending the first stage to help the payload.

It looks like the first stage provided 360 m/s more.  If they use 300 m/s of this for the injection burn (they are not burning to depletion), they could:

(a) go super-sync to an apogee of 67,000 km, reducing deficit to 1630 m/s, or
(b) reduce inclination to about 20 degrees, with GEO apogee, reducing deficit to 1660 m/s or so.

I'd guess they did (b), since that does not need any additional trajectory work by the customer.

It seems you got it right, man! I've got to learn more.......

42070   ECHOSTAR 23   2017-014A     632.93min   22.43deg   35903km X   179km
 

Yes - reduced inclination by 6 degrees ...

[stcks]

It seems you got it right, man! I've got to learn more.......

42070   ECHOSTAR 23   2017-014A     632.93min   22.43deg   35903km X   179km
 

Yes - reduced inclination by 6 degrees ...

I'm calculating GTO-1711. That look right to you all?