Author Topic: Falcon 9 v1.1 - ABS-3A/Eutelsat 115 West B - March 1, 2015 - DISCUSSION  (Read 268016 times)

Offline toruonu

Can someone point to a doc with regardsto how the new electric satellites work with regard to propulsion. Is it a form of ion drive with electricity providing the large momentum for the charged particles or is it using some magnetic fields to adjust itself and accelerate and what kind of limitations that has in usability at random points in the satellites orbits.

http://en.wikipedia.org/wiki/Ion_thruster

Ok, so it's as I expected an ion drive. That does mean that it's still got a limited lifetime depending on the total amount of gas that's packed for the thrusters, but it's nice to see the concept finally making it to real-life usage :) The main reason I was confused was the discussion on Van Allen belts transitions that I somehow remember seeing a bit upthread (or maybe I've been browsing around too much and got confused).

Offline BrianNH

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I can't look at that picture without seeing a gigantic pair of binoculars and wondering if it is secretly an NRO payload.

 :D

Offline Jim

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I expect these commercial electric satellite buses to also be used for planetary missions.

No, not the same environment.  It is a bad idea.  See Mars Observer

Offline Jim

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Ok, so it's as I expected an ion drive. That does mean that it's still got a limited lifetime depending on the total amount of gas that's packed for the thrusters, but it's nice to see the concept finally making it to real-life usage.

This is not the first "real-life usage".  There are many comsats using it.  This will be the first usage for intentional GTO to GSO circularization.

Offline Jim

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So they will go to near GTO and perigee raise/circularize themselves. Like I said, not "LEO."

I mistook their original intentions when they announced these spacecraft a few years ago.

Looks like these spacecraft are going after Orbital's niche.
« Last Edit: 11/13/2014 12:57 pm by Jim »

Offline Norm38

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No, not the same environment.  It is a bad idea.  See Mars Observer

Are you speaking just generally, as in "interplanetary isn't the same as GEO"?  From JPL data, 10 year radiation dose in GEO is 100krad, while in Mars orbit it may only be 5krad.  So anything rad hardened for GEO is good enough for a lot of the solar system.

And Mars Observer failed due to a leaky valve in the hypergolic propulsion system which an all solar-electric craft doesn't have.  But it seems that valve could have leaked just as easily in LEO as on its way to Mars.

Is there something specific that rules the 702SP bus out for wider usage?

Offline Jim

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No, not the same environment.  It is a bad idea.  See Mars Observer

Are you speaking just generally, as in "interplanetary isn't the same as GEO"?  From JPL data, 10 year radiation dose in GEO is 100krad, while in Mars orbit it may only be 5krad.  So anything rad hardened for GEO is good enough for a lot of the solar system.

And Mars Observer failed due to a leaky valve in the hypergolic propulsion system which an all solar-electric craft doesn't have.  But it seems that valve could have leaked just as easily in LEO as on its way to Mars.

Is there something specific that rules the 702SP bus out for wider usage?

It has nothing to do with radiation.  It is the thermal environment and attitude determination. 

GEOsats have one side always facing earth and two sides always facing deep space. 
They also rely on using earth sensors for pointing.
And no, the leak would not have happened on a GEOsat because it would have been used within days of launch.  It wasn't made to keep the tanks isolated that long.

Offline Space Ghost 1962

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It's a question of how much cost/risk you take.

Both in dependence on EP and on reuse for other purposes - e.g. solar system planetary missions.

The paradigm shift underway is highly subjective and driven by volume and market forces, but not commanded by them. What commands is the need to deliver on reliable on orbit service over a reasonable expectation of service life.

At the moment, GSO sat economics are the "life blood" - there's an attempt for EO/LEO to develop speculative products for growth, some which may appropriate/erode significant GSO market as well. But nothing suggests enough cost recovery to qualify such LEO/GSO buses for long cruise, deep space, or other planetary environments in order to address a larger market.

It is a substantial additional cost to do so. When costs are already a pressure point in market growth. To qualify a planetary mission, might be 3x GSO or 8x LEO. Or more.

As to LEO or GEO for EP to GSO - tradeoff is lifetime/risk. Also payload growth margin. Expect that the limiting factor is the number of payloads successfully contracted for and entering service. As experience is gained by vendors/providers, numbers will grow gradually, supporting those providers.

If they over do it, like launching too many, too "low", too short lived ... then we'll find a new bound both on capabilities supplied, as well as a new market "floor" for commercial sat business. If the volume integrated at market price points exceeds same in past, and the supplied services from those sats are cash positive, then the market size may be said to have grown.

Whoever/however that would be accomplished, if it is, would be a historic accomplishment in and of itself.

And that's what the paradigm shift is after, where many have failed. Historically resistant to such growth. 

Offline Space Ghost 1962

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I expect these commercial electric satellite buses to also be used for planetary missions.

No, not the same environment.  It is a bad idea.  See Mars Observer
This bugs me. Reread Mars Observer review board findings.

It was a failure of FBC thinking. Presuming that such a bus could function as such. And intentionally not having the budget to support such additional need, so we could show almost a COTs approach to planetary. Gambling.

It is possible to qualify such busses. This will increase costs greatly. You could improve development, testing, and qualification processes eventually to bring down these additional costs as a structural cost issue. Would it cover all of the necessary requirements of science missions, and the operational requirements to get them there?

None of this would make such busses more commercially viable. All of this would increase cost/risk/time.

So you'd have to "front load" costs, at a time you're shedding them, under the theory of long term gain for infrequently flown planetary missions. There would also be the doubt that it would still be insufficent for science missions - look at the current issues with this comet lander - none of that would remotely be faced by a commercial geosat. We've a long way to go before that, if ever, becomes likely.

add:
But I don't think Mars Observer is a good model for this. As EP upscales, its thermal requirements and other issues for cruise are nothing like the hypers and the antique Tiros bus mentioned. Not to say its a slam dunk like the FBC mentality which I was very wary of at the time. So I think its a bad one to quote because the actual elements are way obsolete.

Big issue for SEP grander use will be PV life/degradation/weight/efficiency. But with more PV usage on earth power generation, there's a large market for PV improvement that is already bringing down PV space issues across the board.
« Last Edit: 11/13/2014 08:02 pm by Space Ghost 1962 »

Offline Jim

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But I don't think Mars Observer is a good model for this. As EP upscales, its thermal requirements and other issues for cruise are nothing like the hypers and the antique Tiros bus mentioned.

It used a RCA Satcom bus, the appendage articulation was from the Tiros bus.

Offline Space Ghost 1962

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But I don't think Mars Observer is a good model for this. As EP upscales, its thermal requirements and other issues for cruise are nothing like the hypers and the antique Tiros bus mentioned.

It used a RCA Satcom bus, the appendage articulation was from the Tiros bus.
If I recall correctly, it was Tiros, then RELAY, then eventually Satcom, then the last series of Tiros was built on the Satcom bus. They were related, have to go back to an octogenarian friend to get his read on this.

All out of David Sarnoff Labs at RCA.

Think the point stands, if not even more reinforced.

Offline Jim

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If I recall correctly, it was Tiros, then RELAY, then eventually Satcom, then the last series of Tiros was built on the Satcom bus.


TIROS as in DMSP and NOAA buses.
http://space.skyrocket.de/doc_sat/lockheed_tiros-n.htm

The MO bus was the  Satcom-K/LM4000 bus
http://www.skyrocket.de/space/doc_sat/lockheed_4000.htm

Online Steven Pietrobon

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Note that the Indians successfully used their I-1K GEO satellite bus for their Mars Orbiter Mission. They learned from the Mars Orbiter and other unsuccessful missions to implement their mission cost effectively in a very short time. The same could probably be done with Boeing's 702SP bus, but it would need to be modified for deep space operations.
« Last Edit: 11/14/2014 03:14 am by Steven Pietrobon »
Akin's Laws of Spacecraft Design #1:  Engineering is done with numbers.  Analysis without numbers is only an opinion.

Offline mr. mark

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I see. Read it wrong. Who writes space articles in Imperial units?  Very annoying.
Spaceflightnow.com is more geared toward a non scientific US average joe audience. Similar to space.com

Offline StarryKnight

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If I recall correctly, it was Tiros, then RELAY, then eventually Satcom, then the last series of Tiros was built on the Satcom bus.


TIROS as in DMSP and NOAA buses.
http://space.skyrocket.de/doc_sat/lockheed_tiros-n.htm

The MO bus was the  Satcom-K/LM4000 bus
http://www.skyrocket.de/space/doc_sat/lockheed_4000.htm
Actually the bus was more like the Series 5000 bus, which had liquid apogee engines (i.e. 100 lbf oxidizer & hydrazine engines) for large orbit adjustments (orbit raising in GTO; Mars Orbit Insertion for Mars Observer).  The Series 4000s had solid fuel Apogee Kick Motors.

Avionics were based TIROS/DMSP buses, which had more autonomy since they were LEOs with just a few ground passes per day. The RCA/GE/MM/LM satcoms of those days didn't have as sufficient autonomy since they were meant to be in continuous ground contact (except short portions during the GTO phase of the mission).
In satellite operations, schedules are governed by the laws of physics and bounded by the limits of technology.

Offline DrLucky

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Hi folks,

I'm trying to educate myself more about electric propulsion spacecraft like these two sats.  This discussion has been pretty informative so far, but I feel like some of it is going over my head.  I thought I'd scrape together what I can find out about the Boeing 702 bus and the propulsion used, and then hope those of you more familiar with the systems to point me to more documents / facts I've missed. 

As always, I welcome correction!

When we speak of "electric propulsion", we generally mean ion propulsion - there's an inert reaction mass electrically propelled at high velocity.  In the case of 702SP, I think it's a(n) XIPS or Xenon-ion propulsion system.

So the Boeing 702 is a family of busses.  I'm assuming that a given bus provides structure, avionics, propulsion, and power in a range of configurations, and the customer then specifies the equipment (reflectors, receivers, etc) that they require.  Anything I'm missing here?

Given the descriptions here: http://www.boeing.com/boeing/defense-space/space/bss/factsheets/702/702SP.page I'm assuming that the 702SP is at the small end of the 702 family (it's extending the range of power down to 3kW)

It seems like the 702SP is unique in the 702 family as being all-electric propulsion?

I tried finding out masses of previous 702 family sats.  I've come up with the Anik F1 (at 3015 kg dry / 4700 kg total) and Anik F2 (at 3489 kg dry / 5950 kg total).  The Galaxies seem similar.  The NSS-8, lost at launch, 5920 kg launch mass, is listed as over 17kW end-of-life power generation, so I'm assuming that's the top end mass of the 702 family bus. (the above article describes 702 as ranging from 3-18kW)

It's harder to get masses on these new sats, but satbeam.com lists the ABS-3A as 1800kg launch mass with no dry mass, and doesn't give a figure for E115WB.

The Boeing document gives "5 kg / year" as the reaction mass consumed per year of operation, presumably for station-keeping.  Given a spacecraft life of 15 years, that means at least 75kg, bare minimum. 

I then dug into what that 5 kg / year buys - this Wikipedia entry http://en.wikipedia.org/wiki/Delta-v_budget gives 50-55 m/s as the yearly delta-v budget for station-keeping in GEO.  I assume this varies by longitude, due to gravitational anomalies, but I wasn't able to find comparison numbers. 

These numbers give very roughly 2000s ISP (one sig fig, at best), which seems to be the right ballpark for ion propulsion.  I've seen "experimental" figures of 210 km/s and 100 km/s thrown about.

Okay, so now, how much propellant (is that the right term?) does it take to get to GEO from LEO? 
I've seen a number of figures for delta-v from an inclined LEO to GEO: The Wikipedia page gives 2 km/s in the "high thrust" section and 6 km/s in the section for low-thrust.  That difference I understand, but then there's a chart farther down which gives 3.8 km/s LEO-GEO.  This seems like a better figure (for actual practice), since I've seen people describing GTO as 1600 m/s from GEO, which this chart echoes.

Using the rough 2000s ISP figure above, that would require about 320 kg (all right - one sig fig? 300kg!) of Xe for the 3.8 km/s insertion, or about 480kg (500) for a 6 km/s, less-efficient burn.

One open question of mine - how many N of thrust are we talking about?  The figure of 4-6 months to enter service has been used, but clearly that's not with a 100% duty cycle on the thrusters, so I've no way to spitball the actual thrust.  Well, except to say that by my (now very crowded) back-of-the-envelope, if you did thrust for 4 months, the 6 km/s burn would be around 1 N.

At any rate, we seem to be talking about an 1800kg spacecraft which is probably less than 500kg xenon, or a 28% PMF.  Anik F2 is about 40% PMF.  It launched on an Ariane 5G+,and was delivered to GTO (according to http://www.spacelaunchreport.com/ariane5.html), so it was required to do a 1600m/s job rather than 3.8 or 6.0 km/s.

So if you can wait 4-6 months, and the sat doesn't get cooked by the radiation belts, it seem like a win.  Are there other considerations I'm missing?  I'm assuming that this doesn't scale up as well, since presumably you'd need outsized solar panels to power larger sats during the thrust phase, which would lead to wasting payload on power that's unneeded on station.  Or will the sat industry just absorb this, make different trades, and put higher-power beams on the sats?

Okay, now for a few other things raised in this thread - XIPS for interplanetary missions was raised, and shot down with Mars Observer as an example.  I'm trying to clearly understand this argument.  My grok was that it wasn't an argument against XIPS itself (it wasn't the electric propulsion which failed on that mission), but rather that a GEO bus (ie: 702SP) hasn't been designed for interplanetary missions, so the services offered by the bus would be at best inappropriate for the mission, or at worst fail in some LOM way because it wasn't designed for the environment?

I recall following Deep Space 1, which was testing a number of techs, including a XIPS called NSTAR.  I believe that engine is also in use on Dawn.  It had 3100s ISP and about 0.1N thrust.  So, presumably, on the right spacecraft, XIPS can be very effective. Or is there info about problems with NSTAR that I'm unaware of?

So, to sum up, I'd love feedback on the following points:
* these spacecraft are about 1800 kg launch mass
* they're perhaps 1300 kg dry mass? (more likely 1600kg+)
* ISP for the thrusters is in the 2000 - 3000s range? (3400-3500s depending on throttle)
* LEO -> GEO for this type of craft is 6 km/s?  (perhaps, but mission is likely GTO->GEO at ~2kps)
* Thrust for the 702SP is a few Newtons? (4 thrusters at 165mN each, possibly operating in concert depending on configuration)
* The argument against 702SP for interplanetary missions is because the bus is inappropriate, not the thruster?

Wouldn't mind hearing what the thruster unit(s) actually are.

Thanks for hearing me out!

Edit: most questions answered.

« Last Edit: 11/17/2014 11:06 pm by DrLucky »
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Offline Kabloona

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DrLucky, here are specs for the 702SP and NSTAR ion thrusters:

http://www2.l-3com.com/eti/product_lines_electric_propulsion.htm

The 25 cm thruster for the 702SP is listed at 3400-3500 seconds Isp and 79-165 mN thrust.
The 702SP carries four of these thrusters according to this source:

http://www.aerospace-technology.com/projects/boeing702/



As for transfer time, according to this article it will take 6-8 months for the transfer from parking orbit to GEO:

http://www.spacenews.com/article/satellite-telecom/39853news-from-satellite-2014-boeing-electric-satellite-backlog-poised-to

Delta V required for this mission was discussed upthread, but some of the posts apparently got lost. To recap, F9 advertised performance to GEO is 4850 kg, and it looks like the dual-payload stack will be in that ballpark, so it's reasonable to expect F9 to put the stack in near-GTO orbit. In that case the satellites will need to supply on the order of 2000 m/sec, certainly not the 6000 m/sec you mentioned for a LEO to GEO transfer.

Given those numbers, I get a ballpark range of 100-150 kg xenon needed for orbit transfer.
« Last Edit: 11/17/2014 05:04 pm by Kabloona »

Offline DrLucky

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Thanks, Kabloona.

I dug up a little more on the NSTAR.  It appears that for DS1, the system dry mass was about 50kg, including 20 kg or so for the propellant feed and tank.  That was for 82kg of propellant.

Interesting that while the NSTAR loses a lot of Isp when operated below max power, the 25cm XIPS can throttle to about 50% without major loss.

That link gives the mass of the 25cm thruster, proper, as 16kg or so, and the power processor (one per thruster pair) as about 21kg. Not sure how the cables and plumbing break down - if that's in addition, for example, but it seems that it's pretty close to being in line with the NSTAR.  Obviously, Xe tankage will be much larger, given the job it has to do.

At 640mN, depending on the plan, the thrusters might need to be on for the full 6 months.

I'm trying to reconcile that with the other stats for the 702SP, though: it advertises 3-8kW of power, yet 4 x 25cm XIPS would be 18kW.  Either I'm misunderstanding something, or they're not running continuously and the bus has substantial energy storage capability.  But batteries are heavy.

With an Isp of 3400s, the propellant estimates I had above are actually high; it's probably more like 300kg of Xe to do the GEO insertion at 6 km/s plus the 75kg stationkeeping budget.


Thanks for the added info!

Edit: apparently I need better reading skills.
« Last Edit: 11/17/2014 07:14 pm by DrLucky »
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Offline Kabloona

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With an Isp of 3400s, the propellant estimates I had above are actually high; it's probably more like 300kg of Xe to do the GEO insertion at 6 km/s

No, read my post again. 6 km/s is from LEO, but this stack will NOT be dropped off in LEO, contrary to the misstatements in this thread.  ::)

It will be dropped in a high elliptical orbit near GTO, if not GTO itself. Thus the transfer delta V required to reach GEO will more likely be around 2 km/s, in which case you're looking at around 100 kg xenon.
« Last Edit: 11/17/2014 05:33 pm by Kabloona »

Offline kevin-rf

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I'm trying to reconcile that with the other stats for the 702SP, though: it advertises 3-8kW of power, yet 4 x 25cm XIPS would be 18kW.  Either I'm misunderstanding something, or they're not running continuously and the bus has substantial energy storage capability.  But batteries are heavy.
More likely, the four thrusters point two or more directions. (Think North, South, East, West for station keeping) So at any one time it is unlikely more than one will be used. 18 kW / 4 = 4.5 kW. That is well with in the 8 kW power budget.

During eclipse season the satellite requires batteries to keep i running at full broadcast power. So it will have large batteries it can tap.
« Last Edit: 11/17/2014 06:23 pm by kevin-rf »
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