SpaceX Falcon Heavy : Arabsat 6A : LC-39A : April 11, 2019 - DISCUSSION

[Alvian@IDN]

Someone can confirm this ? SpaceXNow app state the launch time will be an hour later 

[Crispy]

Might be a mistaken daylight savings correction?

[Alvian@IDN]

Now it's back to normal again

[envy887]

Stephen Clark at Spaceflight Now is reporting that the targeted transfer orbit will have an inclination of 23 degrees and an apogee of 90,000 km. That looks pretty similar to Thaicom 8's transfer orbit, which was about a GTO-1500 (for a 3,100 kg satellite, vs. about 6,000 kg for Arabsat 6A). Do those numbers look about right?

He also gives the satellite mass (6465kg) and says they are using new flexible solar arrays.
https://spaceflightnow.com/2019/04/10/spacexs-falcon-heavy-ready-for-first-commercial-launch/

Using this target orbit and mass, and an assumption of 5,000 kg for the burnout mass of the upper stage, we can calculate the approximate payload to a standard GTO-1800:

A 200x90000 km orbit has a perigee velocity of 10651 m/s. 200 km circular has a velocity of 7784 m/s, including the 5 degree inclination change (28 to 23 degrees) as a vector sum gives a delta-v of 3011 m/s for the GTO burn.

With MVac ISP of 348, the mass ratio for this delta-v is e^(3011/(9.81*348)) = 2.416. So the initial mass in LEO is (6465+5000)*2.416 = 27700 kg.

A standard GTO-1800 is about 2454 m/s from 200 km circular at 28 degrees. This requires e^(2454/(9.81*348)) = 2.052 mass ratio. Assuming the same mass in LEO (which is not a perfect assumption, but should be very close to true), the payload would be 27700/2.052-5000 = 8499 kg or almost exactly 8.5 t.

Repeating with other guesses for the upper stage mass actually doesn't significantly change the baseline estimate for GTO-1800 payload:
3500 kg: 8.23 t
4000 kg: 8.32 t
4500 kg: 8.41 t
5000 kg: 8.50 t
5500 kg: 8.57 t

SpaceX offers 8.0 t to GTO-1800 for $90m. That offered performance appears to be slightly less than the actual performance they intend to give Arabsat regardless the actual upper stage mass.

[gongora]

Repeating with other guesses for the upper stage mass actually doesn't significantly change the baseline estimate for GTO-1800 payload:
3500 kg: 8.23 t
...
5500 kg: 8.57 t

Did you transpose the numbers here?  It looks like you're saying reducing the dry mass of the second stage reduces the performance of the launch vehicle.

[envy887]

Repeating with other guesses for the upper stage mass actually doesn't significantly change the baseline estimate for GTO-1800 payload:
3500 kg: 8.23 t
...
5500 kg: 8.57 t

Did you transpose the numbers here?  It looks like you're saying reducing the dry mass of the second stage reduces the performance of the launch vehicle.

No, the numbers are in the correct order, and no, I am not saying that reducing the dry mass of the upper stage reduces performance.

Recall that I am using a higher energy mission as my baseline. If a heavier upper stage can lift the same mass as a lighter stage to a higher energy trajectory, then it can lift more mass than a lighter stage to a lower energy trajectory.

Another way to look at it is the IMLEO is much higher if the upper stage is heavier, both because the stage is heavier, but also because more prop mass is required for push that heavier stage to GTO. That is, I am assuming the same IMLEO between high energy and low energy GTOs, but I am not assuming that mass in LEO is invariant of the upper stage mass.

[LouScheffer]

Stephen Clark at Spaceflight Now is reporting that the targeted transfer orbit will have an inclination of 23 degrees and an apogee of 90,000 km. That looks pretty similar to Thaicom 8's transfer orbit, which was about a GTO-1500 (for a 3,100 kg satellite, vs. about 6,000 kg for Arabsat 6A). Do those numbers look about right?

He also gives the satellite mass (6465kg) and says they are using new flexible solar arrays.
https://spaceflightnow.com/2019/04/10/spacexs-falcon-heavy-ready-for-first-commercial-launch/

This orbit indeed has about 1500 m/s to go to GEO, and takes about 2950 m/s from LEO.

A canonical GTO (27 degrees, GEO apogee) takes about LEO+2450 m/s.  If you apply the usual second stage numbers, (ISP = 348, fuel= 107t from environmental impact statements, empty + residuals = 5.5t from fitting to LSP numbers) then you find that the second stage could put 8.5t into a nominal transfer orbit.   That's a little more than the 8t SpaceX claims on their web site.

Though we make slightly different assumptions, this is entirely consistent with Envy887s' analysis above.

[HeartofGold2030]

Stephen Clark at Spaceflight Now is reporting that the targeted transfer orbit will have an inclination of 23 degrees and an apogee of 90,000 km. That looks pretty similar to Thaicom 8's transfer orbit, which was about a GTO-1500 (for a 3,100 kg satellite, vs. about 6,000 kg for Arabsat 6A). Do those numbers look about right?

He also gives the satellite mass (6465kg) and says they are using new flexible solar arrays.
https://spaceflightnow.com/2019/04/10/spacexs-falcon-heavy-ready-for-first-commercial-launch/

This orbit indeed has about 1500 m/s to go to GEO, and takes about 2950 m/s from LEO.

A canonical GTO (27 degrees, GEO apogee) takes about LEO+2450 m/s.  If you apply the usual second stage numbers, (ISP = 348, fuel= 107t from environmental impact statements, empty + residuals = 5.5t from fitting to LSP numbers) then you find that the second stage could put 8.5t into a nominal transfer orbit.   That's a little more than the 8t SpaceX claims on their web site.

Maybe the ASDS being further down range than normal (967km vs 600km) contributes to the increased payload, by allowing the core to burn longer?

[intelati]

Maybe the ASDS being further down range than normal (967km vs 600km) contributes to the increased payload, by allowing the core to burn longer?

Hypothetically, yes. I do think we're close to the maximum range with this launch. If you watched the super hot Nusantara Satu landing, you saw the heating sparking the core.

This will be an insane launch honestly.

[TrueBlueWitt]

Stephen Clark at Spaceflight Now is reporting that the targeted transfer orbit will have an inclination of 23 degrees and an apogee of 90,000 km. That looks pretty similar to Thaicom 8's transfer orbit, which was about a GTO-1500 (for a 3,100 kg satellite, vs. about 6,000 kg for Arabsat 6A). Do those numbers look about right?

He also gives the satellite mass (6465kg) and says they are using new flexible solar arrays.
https://spaceflightnow.com/2019/04/10/spacexs-falcon-heavy-ready-for-first-commercial-launch/

This orbit indeed has about 1500 m/s to go to GEO, and takes about 2950 m/s from LEO.

A canonical GTO (27 degrees, GEO apogee) takes about LEO+2450 m/s.  If you apply the usual second stage numbers, (ISP = 348, fuel= 107t from environmental impact statements, empty + residuals = 5.5t from fitting to LSP numbers) then you find that the second stage could put 8.5t into a nominal transfer orbit.   That's a little more than the 8t SpaceX claims on their web site.

Maybe the ASDS being further down range than normal (967km vs 600km) contributes to the increased payload, by allowing the core to burn longer?

I read core is burning >20 seconds longer than last years FH. This is possible thanks to higher thrust and stronger bolted octawebs allowing deep core stage throttling while boosters stay at 100% thrust and do most of the work early.

[LouScheffer]

A 200x90000 km orbit has a perigee velocity of 10651 m/s. 200 km circular has a velocity of 7784 m/s,

So far I follow...

including the 5 degree inclination change (28 to 23 degrees) as a vector sum gives a delta-v of 3011 m/s for the GTO burn.

But here I've lost you.  Calculating the length of the remaining side, using the side-angle-side formula of c = sqrt(a^2+b^2-2*a*b*cos(theta)),
I get sqrt(7784^2+10651^2-2*7784*10651*cos(5*3.14159/180)) = 2975 m/s as the vector sum.

Also, fitting to the LSP escape-trajectory numbers gives an end-of-burn mass (dry stage + residuals) for the second stage as 5.5t.  Very likely this applies to GTO missions as well.  So I think that's the most realistic among your assumed masses.

SpaceX offers 8.0 t to GTO-1800 for $90m. That offered performance appears to be slightly less than the actual performance they intend to give Arabsat

Overall, I completely agree.  FH, with the Arabsat flight profile, should be able to put 8.5t into a traditional GEO-1800 transfer orbit.

[Alexphysics]

Stephen Clark at Spaceflight Now is reporting that the targeted transfer orbit will have an inclination of 23 degrees and an apogee of 90,000 km. That looks pretty similar to Thaicom 8's transfer orbit, which was about a GTO-1500 (for a 3,100 kg satellite, vs. about 6,000 kg for Arabsat 6A). Do those numbers look about right?

He also gives the satellite mass (6465kg) and says they are using new flexible solar arrays.
https://spaceflightnow.com/2019/04/10/spacexs-falcon-heavy-ready-for-first-commercial-launch/

This orbit indeed has about 1500 m/s to go to GEO, and takes about 2950 m/s from LEO.

A canonical GTO (27 degrees, GEO apogee) takes about LEO+2450 m/s.  If you apply the usual second stage numbers, (ISP = 348, fuel= 107t from environmental impact statements, empty + residuals = 5.5t from fitting to LSP numbers) then you find that the second stage could put 8.5t into a nominal transfer orbit.   That's a little more than the 8t SpaceX claims on their web site.

Maybe the ASDS being further down range than normal (967km vs 600km) contributes to the increased payload, by allowing the core to burn longer?

I read core is burning >20 seconds longer than last years FH. This is possible thanks to higher thrust and stronger bolted octawebs allowing deep core stage throttling while boosters stay at 100% thrust and do most of the work early.

Last year's FH also had bolted octawebs, they are needed to integrate the hardware to mate all three boosters, they had to change all the octaweb structure for the side boosters when refurbishing them. The core burns for longer because it can, not because of any additional change. The side boosters won't remain at 100% thrust as they would need to throttle down at Max-Q for the loads on them while going through that part of the ascent. The center core will go for longer just because it can, by throttling down earlier in the launch it reserves more fuel and the side boosters do all the job during that part. The center core from the demo mission just fired for less time because it had to reserve fuel for the long boostback burn and reentry burns it had to perform. This one won't do any boostback burn so it has to reserve fuel just for a longer reentry and for the landing burn. By not spending too much fuel during reentry the center core can go on for a longer period of time.

[TrueBlueWitt]

Stephen Clark at Spaceflight Now is reporting that the targeted transfer orbit will have an inclination of 23 degrees and an apogee of 90,000 km. That looks pretty similar to Thaicom 8's transfer orbit, which was about a GTO-1500 (for a 3,100 kg satellite, vs. about 6,000 kg for Arabsat 6A). Do those numbers look about right?

He also gives the satellite mass (6465kg) and says they are using new flexible solar arrays.
https://spaceflightnow.com/2019/04/10/spacexs-falcon-heavy-ready-for-first-commercial-launch/

This orbit indeed has about 1500 m/s to go to GEO, and takes about 2950 m/s from LEO.

A canonical GTO (27 degrees, GEO apogee) takes about LEO+2450 m/s.  If you apply the usual second stage numbers, (ISP = 348, fuel= 107t from environmental impact statements, empty + residuals = 5.5t from fitting to LSP numbers) then you find that the second stage could put 8.5t into a nominal transfer orbit.   That's a little more than the 8t SpaceX claims on their web site.

Maybe the ASDS being further down range than normal (967km vs 600km) contributes to the increased payload, by allowing the core to burn longer?

I read core is burning >20 seconds longer than last years FH. This is possible thanks to higher thrust and stronger bolted octawebs allowing deep core stage throttling while boosters stay at 100% thrust and do most of the work early.

Last year's FH also had bolted octawebs, they are needed to integrate the hardware to mate all three boosters, they had to change all the octaweb structure for the side boosters when refurbishing them. The core burns for longer because it can, not because of any additional change. The side boosters won't remain at 100% thrust as they would need to throttle down at Max-Q for the loads on them while going through that part of the ascent. The center core will go for longer just because it can, by throttling down earlier in the launch it reserves more fuel and the side boosters do all the job during that part. The center core from the demo mission just fired for less time because it had to reserve fuel for the long boostback burn and reentry burns it had to perform. This one won't do any boostback burn so it has to reserve fuel just for a longer reentry and for the landing burn. By not spending too much fuel during reentry the center core can go on for a longer period of time.

If u want to contradict Elon's tweets? Be my guest. I'm out.

[Alexphysics]

Stephen Clark at Spaceflight Now is reporting that the targeted transfer orbit will have an inclination of 23 degrees and an apogee of 90,000 km. That looks pretty similar to Thaicom 8's transfer orbit, which was about a GTO-1500 (for a 3,100 kg satellite, vs. about 6,000 kg for Arabsat 6A). Do those numbers look about right?

He also gives the satellite mass (6465kg) and says they are using new flexible solar arrays.
https://spaceflightnow.com/2019/04/10/spacexs-falcon-heavy-ready-for-first-commercial-launch/

This orbit indeed has about 1500 m/s to go to GEO, and takes about 2950 m/s from LEO.

A canonical GTO (27 degrees, GEO apogee) takes about LEO+2450 m/s.  If you apply the usual second stage numbers, (ISP = 348, fuel= 107t from environmental impact statements, empty + residuals = 5.5t from fitting to LSP numbers) then you find that the second stage could put 8.5t into a nominal transfer orbit.   That's a little more than the 8t SpaceX claims on their web site.

Maybe the ASDS being further down range than normal (967km vs 600km) contributes to the increased payload, by allowing the core to burn longer?

I read core is burning >20 seconds longer than last years FH. This is possible thanks to higher thrust and stronger bolted octawebs allowing deep core stage throttling while boosters stay at 100% thrust and do most of the work early.

Last year's FH also had bolted octawebs, they are needed to integrate the hardware to mate all three boosters, they had to change all the octaweb structure for the side boosters when refurbishing them. The core burns for longer because it can, not because of any additional change. The side boosters won't remain at 100% thrust as they would need to throttle down at Max-Q for the loads on them while going through that part of the ascent. The center core will go for longer just because it can, by throttling down earlier in the launch it reserves more fuel and the side boosters do all the job during that part. The center core from the demo mission just fired for less time because it had to reserve fuel for the long boostback burn and reentry burns it had to perform. This one won't do any boostback burn so it has to reserve fuel just for a longer reentry and for the landing burn. By not spending too much fuel during reentry the center core can go on for a longer period of time.

If u want to contradict Elon's tweets? Be my guest. I'm out.

Point me to a single tweet where it says the center core wasn't able to keep firing for 20 seconds more after BECO for the demo mission but now it suddenly can or one where he says that the boosters would run at full thrust all the time and the center would do all the throttling.

[envy887]

A 200x90000 km orbit has a perigee velocity of 10651 m/s. 200 km circular has a velocity of 7784 m/s,

So far I follow...

including the 5 degree inclination change (28 to 23 degrees) as a vector sum gives a delta-v of 3011 m/s for the GTO burn.

But here I've lost you.  Calculating the length of the remaining side, using the side-angle-side formula of c = sqrt(a^2+b^2-2*a*b*cos(theta)),
I get sqrt(7784^2+10651^2-2*7784*10651*cos(5*3.14159/180)) = 2975 m/s as the vector sum.

Also, fitting to the LSP escape-trajectory numbers gives an end-of-burn mass (dry stage + residuals) for the second stage as 5.5t.  Very likely this applies to GTO missions as well.  So I think that's the most realistic among your assumed masses.

SpaceX offers 8.0 t to GTO-1800 for $90m. That offered performance appears to be slightly less than the actual performance they intend to give Arabsat

Overall, I completely agree.  FH, with the Arabsat flight profile, should be able to put 8.5t into a traditional GEO-1800 transfer orbit.

I used sqrt((10651*sin(5 degrees))^2+(10651-7784)^2) but you're right that this isn't the correct formula. Serves me right for trying to derive trig formulae in my head. However, the slightly lower delta-v doesn't materially change the answers I get.

[Tass]

For some reason the webcast on spacex.com/webcast is set to start in seven hours (1:35 UTC 3:35 for me), I thought it was supposed to be 18:35 EDT, 22:35 UTC. Does anybody know anything? Is it just a mistake?

I am willing to wait up 'til half one AM, but not four.

[d3jf]

For some reason the webcast on spacex.com/webcast is set to start in seven hours (1:35 UTC 3:35 for me), I thought it was supposed to be 18:35 EDT, 22:35 UTC. Does anybody know anything? Is it just a mistake?

I am willing to wait up 'til half one AM, but not four.

That’s a mistake. Launch still on for 22:36 UTC.


Sent from my iPhone using Tapatalk

[I14R10]

I saw that too. It's now "live in 3 hours"

[intelati]

Should be home and changed right before the final countdown. Keeping my eye peeled for any twitter updates

Here's to a smooth flight!

[seawolfe]

this is so cool to have NSF commentary and video channel along with the forum!