SpaceX FH : Falcon Heavy Demo : Feb 6, 2018 : Discussion Thread 2

[Herb Schaltegger]

Does anyone know if there will be a parking orbit with a later "TMI" burn ... or will it be a direct ascent with a single S2 burn.

I was hoping somebody actually knowledgeable would answer this. It was discussed before and I believe the best guess was direct.

I don't think that accurately represents the current best guess.  I thought the most convincing argument was that they would want to launch at a time such that all important flight events occurred in sunlight.  i.e. That way all the separations, etc. would have good recording instrument visibility to allow post-mission analysis (or failure analysis if necessary). But the timing required to get to the proper orbit would mean they were launching into darkness if they tried for a direct injection.  So, it seems like they'll want to launch earlier with at least a short section of parking orbit prior to TMI.  This would also give them the maximum width of launch window and let them hold some if needed.

Since they're not actually aiming to intercept Mars, the single largest constraint to their window - besides Range availability and sub-cooled LOX-related technical stuff - is frankly not hitting the Moon. They can launch direct easily enough, which might in fact explain the apparently-lofted trajectory evidenced by the recovery barge position on this mission.

[OneSpeed]

What is the hypothetical trajectories for a Falcon Heavy Demo Mission which will launch at 1:30 EST on February 6th 2018?

Launch at 13:30 EST rules out direct injection to Mars heliocentric orbit. However, the initial launch trajectory might still be quite lofted, and launch anywhere in the 13:30-16:30 EST window should be feasible. A launch at say 3pm Eastern time would be oriented as in the first image, and the burn to an elliptical LEO could be as in the second image. The TMI burn would occur above the terminator about 18 minutes later (depending on the exact launch time). The transit to Mars orbit would take about 160 days.

[Jdeshetler]

Approx, how high is FH by the time it clear the perimeter fence in 30 seconds?

[OneSpeed]

Approx, how high is FH by the time it clear the perimeter fence in 30 seconds?

If the trajectory really is this lofted, the instantaneous impact point (IIP) would reach the perimeter about 26 seconds after launch, at an altitude of about 2,000m. The rocket itself would clear the perimeter in about 32 seconds, at an altitude of about 2,900m.

[IanThePineapple]

Approx, how high is FH by the time it clear the perimeter fence in 30 seconds?

If the trajectory really is this lofted, the instantaneous impact point (IIP) would reach the perimeter about 26 seconds after launch, at an altitude of about 2,000m. The rocket itself would clear the perimeter in about 32 seconds, at an altitude of about 2,900m.

So really after 30-40 seconds we can stop worrying about the pad getting severely damaged/destroyed 

[deruch]

I don't think that accurately represents the current best guess.  I thought the most convincing argument was that they would want to launch at a time such that all important flight events occurred in sunlight.  i.e. That way all the separations, etc. would have good recording instrument visibility to allow post-mission analysis (or failure analysis if necessary). But the timing required to get to the proper orbit would mean they were launching into darkness if they tried for a direct injection.  So, it seems like they'll want to launch earlier with at least a short section of parking orbit prior to TMI.  This would also give them the maximum width of launch window and let them hold some if needed.

Since they're not actually aiming to intercept Mars, the single largest constraint to their window - besides Range availability and sub-cooled LOX-related technical stuff - is frankly not hitting the Moon. They can launch direct easily enough, which might in fact explain the apparently-lofted trajectory evidenced by the recovery barge position on this mission.

My understanding was a recap of Lou's comment with some subsequent discussion by others: https://forum.nasaspaceflight.com/index.php?topic=42705.msg1767687#msg1767687

Basically, that you want your departure burn to end pretty close to the terminator (~1800hrs), so that with Earth's gravity turning you you depart at ~midnight.  But, if SpaceX waited until the end of their upper stage, direct injection burn would coincide with the terminator, that would mean launching late enough that there would be reduced illumination and therefore impaired visibility of the separation events.  As this is a test launch, being able to do post flight analysis of these events is of high value.  Hence the belief that they would launch earlier and not do a direct injection. 

Given the announced launch windows, I believe this has been borne out.

[Jdeshetler]

See bold...

Holding my breath for up to 30-40 seconds? That's doable....

So really after 30-40 seconds we can stop worrying about the pad and NSF's remote cameras getting severely damaged/destroyed 

[Lar]

++

Please remember that when you write about reporters in this industry, friends and colleagues may be reading.

Yes. This is a special place in part because we don't do that sort of thing. Just not done.

Except for Andy Pasztor?

(fan)Not a reporter. Writing dreck for the WSJ doesn't make him a reporter.

(mod) Your point is valid, we probably could do better there.

[kerrycockram]

See bold...

Holding my breath for up to 30-40 seconds? That's doable....

So really after 30-40 seconds we can stop worrying about the pad and NSF's remote cameras getting severely damaged/destroyed 

I have a spare Canon 7D2 with a 10-22mm lens I'd be willing to risk. Sorta kinda kidding, but mostly kinda not

[Roy_H]

What is the hypothetical trajectories for a Falcon Heavy Demo Mission which will launch at 1:30 EST on February 6th 2018?

Launch at 13:30 EST rules out direct injection to Mars heliocentric orbit. However, the initial launch trajectory might still be quite lofted, and launch anywhere in the 13:30-16:30 EST window should be feasible. A launch at say 3pm Eastern time would be oriented as in the first image, and the burn to an elliptical LEO could be as in the second image. The TMI burn would occur above the terminator about 18 minutes later (depending on the exact launch time). The transit to Mars orbit would take about 160 days.

I'm having trouble making sense out of the last picture showing the orbits of the inner planets and Falcon/Roadster. I see small white dot, small red dot and large dot on each orbit. I assume these dots represent position at the same time. I also assume that the planetary motion is counter-clockwise. I would have thought that one of the dots for Earth and Falcon would be coincidental (at launch) but this is not the case. First are my assumptions correct, and second could you identify times to the dots, and third explain why there is no coincidental dot at the point where Earth and Falcon orbits meet. Thanks.

Edit: so I see red dots and white dots are 180° apart and have little or nothing to do with time. Looks like apogee & perigee.

[OneSpeed]

I'm having trouble making sense out of the last picture showing the orbits of the inner planets and Falcon/Roadster. I see small white dot, small red dot and large dot on each orbit. I assume these dots represent position at the same time. I also assume that the planetary motion is counter-clockwise. I would have thought that one of the dots for Earth and Falcon would be coincidental (at launch) but this is not the case. First are my assumptions correct, and second could you identify times to the dots, and third explain why there is no coincidental dot at the point where Earth and Falcon orbits meet. Thanks.

Edit: so I see red dots and white dots are 180° apart and have little or nothing to do with time. Looks like apogee & perigee.

Yes, motion is counter clockwise, the red and yellow dots are apogee and perigee, and the time is launch + 160 days. The white dots are the roadster's trajectory.

[Roy_H]

Thanks for the clarification. Now a new one, out of curiosity. All transfer orbits to Mars that I see are Hohmann and about 200+ days. But Elon talks about 77 day trips. Would these 77 day trips come up once every two years and when?

[speedevil]

If the trajectory really is this lofted, the instantaneous impact point (IIP) would reach the perimeter about 26 seconds after launch, at an altitude of about 2,000m. The rocket itself would clear the perimeter in about 32 seconds, at an altitude of about 2,900m.

Though they could cant the trajectory over a little for the first ten seconds, and get out moderately earlier with essentially no penalty.

Once the vehicle passes perhaps a kilometer or so in altitude, ~18s in, all that hits is a confetti of tank, and all of the lumps, with no majorly significant flame reaching the ground, and no big explosions to damage stuff.
Just the octoweb/engines, fairing, and helium tanks..



This is a not particularly relevant grasshopper launch leading to termination at about a kilometer up.
It had a very small percentage of nominal fuel, and only one core not three, of course, and seemed to be slowing down.

[OneSpeed]

Thanks for the clarification. Now a new one, out of curiosity. All transfer orbits to Mars that I see are Hohmann and about 200+ days. But Elon talks about 77 day trips. Would these 77 day trips come up once every two years and when?

It all comes down to how much ∆V you have available. Taking the Roadster as an example, the sim predicts that there is enough ∆V for a heliocentric orbit with an aphelion of 1.58 AU. If I reduce the S2 burn by 6 seconds, the aphelion reduces to 1.42 AU, and the time taken to reach Mars orbit increases to about 210 days (shown below). Conversely, if the FH had more performance (e.g. block 5), the aphelion could be increased, and the transit time reduced.

Of course the roadster is not actually going to Mars, but a 77 day transit is feasible at any conjunction given sufficient ∆V from LEO (or MEO). This is why it is so important that BFR will support in orbit refuelling. It will enable missions unthinkable with a single-shot architecture.

[.Scott]

Launch at 13:30 EST rules out direct injection to Mars heliocentric orbit.

Is that really true?  I believe that a direct injection could result in a Mars heliocentric orbit over a wide launch window - including 3pm.  But if you were looking to put the Earth orbit at the perigee, that would restrict the window.

[Jim]

Launch at 13:30 EST rules out direct injection to Mars heliocentric orbit.

Huh?  There is no planetary target and so the launch window can be anytime.

[LouScheffer]

Launch at 13:30 EST rules out direct injection to Mars heliocentric orbit.

Huh?  There is no planetary target and so the launch window can be anytime.

Depends on the object of your demo.   If your target is to get an orbit with apogee at Mars distance, you need to direct inject at about 18:00.  Otherwise you are not getting the full benefit of the Earth's speed around the sun.  You'll still get a heliocentric orbit, but the apogee will be short of Mars orbit.

[.Scott]

Launch at 13:30 EST rules out direct injection to Mars heliocentric orbit.

Huh?  There is no planetary target and so the launch window can be anytime.

Depends on the object of your demo.   If your target is to get an orbit with apogee at Mars distance, you need to direct inject at about 18:00.  Otherwise you are not getting the full benefit of the Earth's speed around the sun.  You'll still get a heliocentric orbit, but the apogee will be short of Mars orbit.

I am not sure.  Let me ask the question this way: If you are putting a payload into heliocentric orbit from Earth with the apogee at the Mars orbit, would it take more energy to put it there with an Earth-orbit perigee or a Venus-orbit perigee?

I don't know the answer.  But without considering the lower perigee, you cannot conclude that direct injection at 16:00 would not reach a Mars-orbit apogee with the same or less propulsion as the 18:00 injection.

[Ben the Space Brit]

A question that just occurred to me: I think that we're all in agreement that the initial orbit of the Tesla Roadster will take it nowhere near Mars due to a difference in alignment. Is there any likelihood of precession or orbital mechanics bringing the Roadster to Mars in less than 100 years?

[Kabloona]

If you are putting a payload into heliocentric orbit from Earth with the apogee at the Mars orbit, would it take more energy to put it there with an Earth-orbit perigee or a Venus-orbit perigee?

Since you're starting from Earth, the minimum-delta-V heliocentric orbit with an apogee at Mars has its perigee at Earth. To reduce the perigee from Earth to Venus, the minimum delta-V solution would require a retrograde burn at apogee, ie more energy.