NASA’s Selection of the Blue Moon Lander for Artemis V

Posted on May 25, 2023 by Jonathan Goff

Last week, when NASA announced that they were picking Blue Origin’s National Team to develop a sustainable human lander for the Artemis V mission, what surprised me wasn’t the selection, but the fact that I’ve come around to really liking the decision.

NASA’s Associate Admin for Exploration Systems Development, Jim Free, at the Artemis V lander selection press conference (Credit: NASA/Aubrey Gemignani)

While it’s still recent enough to be relevant, I wanted to share some thoughts on Blue Origin’s updated human lander architecture, why I think this was the right selection in spite of my feelings about their original concept, thoughts on the execution challenges they’ll face, and some of the interesting future possibilities having two fully-reusable lander architectures may open up for NASA. But first, you may be wondering why I was so surprised that I would end up liking Blue Origin’s lander architecture.

Why I Wasn’t A Fan of Blue Moon 1.01

I hope I don’t offend any of my friends who work or worked at Blue Origin by saying this, but if you had to summarize my initial reaction to the National Team’s original HLS lander concept, I would’ve used the word cynical.

It felt cynical, because rather than trying to come up with the best solution for affordably and reliably bring people to/from the lunar surface, they seemed instead to be regurgitating exactly what they thought NASA wanted to see. NASA had a reference mission concept that was a complex, fully expendable system with three stages — an in-space tug, a descent element, and a separate ascent element, so that’s what they had. Blue Origin was a relatively unproven space contractor, so they added not one, but two aerospace primes to their team. Which also happened to maximize the number of congressional districts their project would have work performed in2. It almost felt like they weren’t even trying to win, so much as guarantee that they’d be one of the two solutions picked3. While in theory at least some of the elements in Blue Moon 1.0 could be refueled and reused, some components like the massive lander descent stage had no easy path to future reuse.

The National Team’s original Blue Moon HLS concept, as proposed in 2020 (Credit: Blue Origin)

In contrast, the Dynetics ALPACA design was a creative approach that seemed to be genuinely trying to provide NASA with a good way of getting people and cargo to and from the lunar surface. By the end of the base period, they had shifted to a single-stage lunar lander concept that leveraged in-space refueling, and had a clear pathway to full reuse. The low-slung central crew/cargo attachment point allowed easily delivering crew to/from the lunar surface as well as delivering large cargo modules, without needing multi-story ladders or elevators. The low CG meant that it could likely land on rough terrain with a lower odds of tipping than a design like Starship HLS. It did have the teensy problem that at the time proposals had to be submitted, the design’s mass budget didn’t close yet4, but they did eventually close the design, just not in time for consideration in the Option A evaluation.

Autonomous Logistics Platform for All-Moon Cargo Access (ALPACA) lander concept from 2020 (Credit: Dynetics)

Anyhow, suffice it to say, that going into last week’s HLS Sustainable Lander Development announcement, I was really rooting for Dynetics to win, and didn’t have a very high opinion of Blue Origin’s lander concept. So, when Administrator Nelson announced the National Team had been selected, my first reaction was pretty strong disappointment.

I’m glad I was too busy taking notes from work to tweet my immediate hot take, because by the end of the call, my opinion had shifted pretty dramatically.

Why Blue Moon 2.05 is a Dramatic Improvement Over v1.0

It took the press conference a while to show any details about the new design, but when they unveiled the Blue Moon 2.0 lander, I almost did a double take. The design was superficially similar to the original design, but you pretty quickly noticed some pretty significant differences. Was I seriously seeing a bottom-loader single stage design?

Meet the Blue Moon 2.0 lander concept (Credit Blue Origin)

I never got around to blogging about what I call bottom-loader SSTO landers, but it’s an idea I first learned about almost 20 years ago with t/Space’s CXV Stage 2 concept from their Concept Exploration & Refinement study final report6. It’s one of my three favorite Unorthodox Reusable Lunar Lander Concepts that I’ll hopefully get more chance to blog about in the future. Needless to say, when I saw that, I perked up and started paying real attention.

If I had to summarize the highlights we could glean of the proposed Blue Moon 2.0 lander architecture, I’d point out three key features:

  • Bottom-Loader SSTO Lander: Crew or cargo pod on the bottom, propellant tanks on top. Enables easy surface access without cranes or ladders. Keeps the CG low for reduced tipping risks. Keeps the load paths for the rocket higher efficiency. Provides the best thermal isolation between the warm parts7 and the parts that want to be kept really, really cold8. This is the part that Blue Origin would be developing
  • Reusable Cis-Lunar Refueling Tug: While they never showed any pictures of the tug, this element would bring LOX and LH2 propellant from LEO to NRHO to refuel the lander, and return to be refueled again for reuse. This is the part that Lockheed Martin would be developing.
  • Reusable From the Start and ISRU Compatible: By going to a single-stage architecture, there’s a clear and easy path forward to refueling — initially in NRHO using tugs coming from LEO, but eventually also on the lunar surface9. Also, while LH2 is harder to handle than Methane, LOX/LH2 can be derived readily from lunar water ice sources10, enabling a switch over from terrestrial to lunar sourcing once ISRU is proven out/debugged/scaled up.

In short, Blue Origin responded to their Option A loss in 2021 by significantly improving their offering to NASA, offering a solution that was innovative and actually worth funding.

Why I think Blue Moon 2.0 Was the Right Call

I’m still a fan of Dynetics’ ALPACA and LLAMA concepts, and I hope they find some way to see the light of day. But given what we know now, I think the Blue Moon 2.0 concept was the right call for NASA, and not a politically-motivated decision, or one that only won because a space billionaire bought his way to success.

First, and most importantly, I think Blue Moon 2.0 helped close the innovation gap between the National Team and Dynetics. Blue Moon 2.0 captures many of the benefits that ALPACA brought to the table, including: lower CG for better landing on uneven terrain, crew/cargo located close to the ground for easy ingress/egress and loading/unloading, ability to deliver significant cargo mass to the surface, and a single-stage design with a clear path to full-reusability. In some ways it was better than ALPACA, by providing a cleaner load path and easier thermal isolation of cryogenic tanks from heat sources, a propellant combo that had an easier path to 100% sourcing from lunar ISRU, and a more developed fully-reusable cislunar tanker concept11. There were some relative drawbacks like the challenges of LH2 storage, and the potentially smaller available cargo volume12, but overall they did a good job of narrowing or closing the gap with Dynetics’ solution.

Second, there seemed to be far less zip-code engineering this time around13. There are multiple team members still, but each of them makes logical sense, not just as a way to get more congressional support.

Third, while there’s very real execution risk for Blue, since they haven’t flown anything anywhere near this complex, Dynetics carries similar risk, so it’s not really a discriminator.

Fourth, while Bezos’s willingness to subsidize the price to NASA probably made a difference, it was far from the only consideration, and in my opinion was probably more of icing on the cake. In addition to closing the innovation gap with Dynetics, it sounds like Blue did a better job of convincing NASA that they had a design that unambiguously closed technically. If Dynetics had still had the clearly superior concept, and if they had done a better job of making it unambiguous that they had a design that closed for all of NASA’s needs, I think they would’ve had a decent chance of winning, even with being more expensive to NASA.

In the end, for all of these reasons, I think NASA made the right call. That said, while I doubt it will happen, I hope Dynetics finds some way to get their concept fielded14.

But Can Blue Deliver?

This is where I have the strongest reservations. While Blue has now laid out an innovative and exciting architecture that’s worthy of being funded, a concept is only as good as the organization tasked with executing it. And frankly, people have reasons to have reservations about Blue Origin’s ability to execute on a project this complex. Whether you look at how late the BE-4 engines were in development, how long it took New Shepard to transition from flight test into operations, or how long New Glenn has been taking to make visible progress, there’s definitely room to worry that Blue sometimes take the Graditim part of its slogan more seriously than the Ferociter part. One thing Blue Origin has made me realize is that while I’ve had too much experience with having too little money, that there are real risks in having too much money, that has too few requirements for demonstrated traction tied to it.

It’s an open question if Blue can change its company culture and processes quickly enough to be able to deliver on an ambitious project like this on a tight schedule. I hope they can succeed at that evolution though, because if both the National Team and SpaceX are successful, it could lead to a very exciting new world.

What If They Are Successful?

If both SpaceX and the Blue Origin National Team are successful, we enter a really interesting world. As Eric Berger pointed out in this Ars Technica article today that I was quoted in, both architectures are now based solidly on the use of reusable launch, in-space cryogenic storage and transfer, and in-space reuse. As I pointed out in the intro to my unfinished series on Unorthodox Reusable Lunar Lander Concepts, a fully-reusable lander architecture brings a lot of advantages, beyond the obvious ones of cost savings:

  • Lower Marginal Costs: While you’ll still have some fixed costs associated with the lander infrastructure, the marginal cost of such an architecture drops dramatically, since you’re not having to build new lander or in-space tug hardware for every mission.
  • Throttleability: Once you have a stable of multiple reusable landers, where there aren’t any major expendable components, it becomes a lot easier to throttle up or down mission tempos based on budget availability. If you have a year or two that you need more money to fund say Mars system development, you can throttle down to a lower ops tempo without risking losing the capability, unlike what happened during Apollo.
  • Easier International Involvement: While Starship and New Glenn should theoretically be cheaper than any other launch source, if NASA is paying for those launches, it’s still a cost. But with a distributed lift/tanker architecture, it becomes more feasible to allow international partners to contribute propellant or crew or cargo launches to LEO as their part of the mission. Even if their rockets are more expensive, if NASA isn’t having to pay for those launches, it lowers the cost to NASA.

In addition to those benefits, a fully-reusable Cislunar tug, like what LM is proposing as their part in the Blue Moon 2.0 architecture, opens up some very interesting possibilities. Once you have a reliable way of getting from LEO to NRHO and back reusably with propellant, it’s a relatively straightforward upgrade to add the ability to ferry crew and/or cargo instead of or in addition to propellant. And since Blue will have already developed a crew cabin that’s safe for up to 30 days on the lunar surface, using a derivative of that as a crew pod on the reusable Cislunar Tug isn’t a crazy option. We don’t have hardly any details on LM’s concept, so there’s a chance they might have something in mind that wouldn’t be able to do the LEO-NRHO-LEO loop with crew or cargo, but most of the most likely options should be fairly straightforward to do that.

Once you have the ability to move crew, cargo, and propellants around from LEO to NRHO with a fully reusable system, do you really need SLS and Orion anymore? The vast majority of the budget being attributed to Artemis was the development and operation of SLS and Orion, but they’re only really capable of one mission per year. If you replaced them with distributed lift and reusable Cislunar tugs for crew/cargo out to NRHO, you could probably enable upping the lunar mission tempo dramatically, while freeing up money for developing lunar surface habitation and ISRU payloads. It’s still a longshot politically, but if SpaceX and the National Team are successful, we could be living in very interesting times.

Posted in Blue Origin, Commercial Space, Lunar Exploration and Development, Propellant Depots, Reusable Lunar Landers, Space Transportation, SpaceX | Tagged , , , , , | 7 Comments

Regenerative cooled turbopump

Posted on March 16, 2023 by johnhare

This started as a short thought on nasaspaceflight and grew into something that resembled a blog post. So I copied and pasted it here. I screwed up the link but this is an ongoing thought from the turbine in chamber stuff from a decade and a half back. November 10 2008 Performance Monoprop being one of the series. I recently realized that the concept could serve as a pressure booster in a gas generator cycle to get performance almost matching that of staged combustion with a higher thrust to weight ratio and much faster and cheaper development.

I seem to be drifting back to an ability to do some simple prototyping and am interested in finding an engine company that might do a bit of business with an inventor.

The provenance of the regenerative cooled turbopump includes Rotary Rocket, ATREX, and a patented “single rotor turbine” from LANL. https://image-ppubs.uspto.gov/dirsearch-public/print/downloadPdf/6807802

Rotary Rocket introduced a concept to avoid turbopumps (kinda) by putting the thrust chambers on the ends of rotating arms. The canted thrust chambers spun the assembly which pressurized the propellants in the feed pipes with the centrifugal force. There was no central drive shaft as the assembly freewheeled. It eliminated the separate gas generator, turbine, drive shaft, impeller, and all of their housings and accessories. The engine was never completed beyond some subscale testing. Which makes sense when one thinks of the problems of getting 72 thrust chambers to operate properly in the high gee field at the ends of those arms.  Just the weight of all those engines rotating at high speeds should make one nervous. Still, the idea of simplifying the engine process is appealing.

The ATREX was a Japanese hydrogen fueled air-turborocket with tip turbine blades mounted to the outside of the compressor. It was claimed to be lightweight and compact. Whether all the claimes made sense or not I don’t know. But putting tip turbine blades on the Roton structure seemed to be an easier path than the mass of  whirling engines. It could be a much smaller structure delivering propellants into a thrust chamber bolted onto the vehicle in the normal way. Then the gas  in the thrust chamber could drive the blades on the way through the throat. Seemed simpler, but exposes the turbine blades to more heat than even regeneratively cooled blades would be able to stand. Also, it leaves no path for cooling the nozzle and thrust chamber.

The Single Rotor Turbine https://image-ppubs.uspto.gov/dirsearch-public/print/downloadPdf/6807802 patented by LANL has the hollow turbine blades as the last stage of the compressor. The air flows through the inside of the blades as coolant while being accelerated and compressed. Then the air flows out of the blades into a volute for pressure recovery before entering the combustion chamber. In this way, all of the air is used as turbine coolant before all of it is burned in the combustion chamber and used to drive the turbine. The goal being to raise the allowable turbine inlet temperature for a more efficient engine. Primary direction seemed  to be power generation. There seems to quite a bit of prior work with ideas of this nature as several previous patents are referenced.

Liquid propellants by nature have about three orders of magnitude more mass for a given volume than gas. As such, there is an enormous amount of coolant available inside the turbine blades limited by the requirement that the propellants exit the turbine blades as liquid. It also takes far less energy and velocity to pressurize a given amount of liquid than it does to pressurize gas. It seems apparent that using the LANL concept applied to liquid propellant would result in a relatively simple high pressure turbopump.

In the last post on the regenertive cooled turbine, I skipped over the pump characteristics suggesting using it in a fairly normal layout that happened to allow much higher turbine inlet temperatures. This actually is better as a stand alone turbopump as the driveshaft, housings, and torques involved would be just as heavy as the standard units for a fairly modest gain in capabilities at the expense of R and D on a new system, not to mention the uncertainty. Also, it would almost certainly be tasked to a staged combustion engine which is one of the most expensive and difficult engines to develop.

What I am going to suggest here is more modest in some ways and more radical in others. Develop a small turbopump and gas generator of this nature that free wheels on its’ shaft in the same manner as the Roton engine. Without having the stresses of torque through the turbine blades and disk, torsion in the shaft, and torque limitations in the impellers, the single rotor turbopump can operate at much higher speeds than any normal layout. The turbine inlet temperature can be much higher than any normal turbine even while the disk and blades are much cooler. The torque stresses are minimal with the drive gas on one side of the turbine blade driving against the liquid propellants on the other side of one thickness of metal. The centrifugal stresses will be the same for a given rpm and radius.

With the “brakes off” as compared to normal systems, the tip velocities of the turbine/impeller combo can reach speeds normally reserved for turbo compressors. Tip velocities creating velocity head, and velocity head going as the square of velocity, 40,000 feet of head pressure is theoretically achievable with tip velocities of 1,600 fps.  LOX to over 19,000 psi and RP to over 14,000 psi is theoretically achievable. Cutting those in half with decent pressure recovery might give something close to reality. Using the RP at 7,000 psi as pressure in the gas generator and running as hot as the turbine blades allow should bring the turbine pressure drop to a very modest value allowing 5,000+ psi at turbine exit.

The interesting thing about the single rotor turbopump is that it doesn’t gain weight at the same rate as a conventional system. One that would run all the RP and much of the LOX, or vice versa, in a Merlin would be a unit you could pick up in one hand and not strain. But that would be several bridges too far. Especially as on of the weaknesses of this system is that it gives no reasonable path for regenerative cooling of the thrust chamber and nozzle.

I suggest it might be worthwhile to use such a unit in place of the gas generator on a gas generator cycle engine. Gas generators seem out of style at the moment with so many going for the staged combustion and full flow staged combustion. The reasons are the improved performance compared to the gas generator due to the gas generator exhaust being both lower temperature and lower velocity than the main flow. That percentage of propellant “not pulling its’ weight” is a potent argument. I suggest that replacing the passive gas generator (illustration 2-20 Huzel and Huang page 45) with a very small active gas generator with the single rotor turbopump could boost gas pressure to the standard turbine by 5,000 psi. This could allow either much less propellant to run the standard turbine which would boost system Isp. Or it could have and exhaust of 1,000 psi into a secondary combustion chamber with additional oxidizer creating a respectable Isp in itself. If the gas generator is using 10% of the total and getting 320 seconds in vacuum and the main chamber is getting 360, then the system would only be about 3-4 seconds below that obtained with staged combustion at similar pressures. Considering that the gas generator cycle has inherently better thrust to weight and is easier to develop, a new conservatively designed gas generator engine might be a contender against the state of the art stuff being developed now.

And then there is the option of using them as boost injectors. Initially one. Then later possibly 19 in the manner that the V2/A4 engine recycled the burner cups from a previous.

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Second Guessing Starship

Posted on September 19, 2022 by johnhare

There seem to be two extremes when it comes to Elon Musk and SpaceX. There are those that believe everything Elon tries will work because–ELON. There are those that believe SpaceX has reached the end of its’ lucky tether because–ELON. The extreme factions make it a bit more difficult for those of us that try to think, or second guess , what will unfold in the SpaceX universe.

Starship typifies the extremes. Some believe that Starship will be in full service within a year delivering 100+ ton payloads multiple times per day at a cost per flight well under even Falcon9. There are others that don’t see Starship becoming operational in this decade. Again the extreme factions throw a bit more murk out there to peer through.

There are still quite a few of us somewhere in the middle that doubt both extremes. For my part, I suspect it will become operational mid decade at a price point above Falcon9 per launch, though well under it per unit of mass. For heavy lift, SLS isn’t even on the radar of those that wish to open up space as a place for humanity to live, work, and grow.

Where I second guess Starship and Superheavy is on the sequence used in the development path. I felt there was a good bit of hubris in going for the largest launch vehicle in history with several concurrent new technologies in one jump. I did not strongly hold these opinions until seeing the difficulties with ground support equipment, regulatory approvals, and ersatz-environmentalist tantrums. Hence Second Guessing. This is not a recommendation going forward, rather a hindsight thought.

Considering that many of the issues are caused by the sheer size of the vehicle, it is worth considering what could have been done with a more modest precursor vehicle. For this post, I am suggesting a Falcon9 type layout, though that may not be the best possible size or configuration. 9 Raptor engines first stage and one Raptor vac on the expended second stage.

(simplifying) Assuming that Raptor has twice the thrust and a higher Isp than the Merlin, the Falcon/Raptor (FR) should have a bit over twice the payload of the Falcon9. Assuming that Elon and company have accurately forecast that Starship will cost less per launch than Falcon, then the configuration here should have a launch cost about 25% that of Falcon9. Assuming that the 3 launches per day per Superheavy are not a total fantasy, then the FR booster should be able to hit a cadence of once per day with RTLS.

Assuming that had been done, what would it have gained? For starters, the Starship test flights last year would have been first stage FR test flights. Those test flights could have been followed by full up test flights by the end of last year with operational missions starting early this year. By now it seems reasonable that launch cadence could be approaching that of the Falcon9.

Construction of (hand waving) 5 meter diameter first stages would take about half the time and materials per unit of height as either Starship or the Superheavy boosters. Remember the problems getting the welds right early on? Learning curve (time) could possibly been halved. Along with smaller faster ground equipment construction with about a quarter of the propellant volume requirements per test. Problems could have been found and addressed even sooner than the current gargantuan effort.

Regulators seem to be having problems with the size of the Superheavy Starship combo. Kilotons of equivalent explosive seems to come up on a regular basis. Decibels created by the largest launch vehicle in history gets some of them going. It seems possible that a FR could be passed off as the next logical Falcon upgrade. Somewhat larger to accommodate the more environmentally friendly fuels. A diameter and height that could (theoretically) use the same launch pads as Falcon9. The higher payload is just a byproduct of more efficient engines, nothing to worry about. Pads at the cape and Vandenberg could be modified relatively quickly to handle both types of Falcon

Environmentalists, real and ersatz, would have a much harder time fighting a vehicle that was just a bit larger than what was in the original EIS, really just a minor upgrade in the scheme of things. With no immediate need for massive expansion due to the smaller vehicle, there would have been far less they could try to block. Everything necessary being permitted and underway before drawing their serious attention.

If the Superheavy/Starship combo hits full operational status shortly, then most everything in this post can be used to ridicule me. If the costs per launch are in the single digit millions within a couple of years, the ridicule can be redoubled. I would accept that ridicule with a smile, and possibly a few belly laughs. I do not expect it to just coast to a smooth operational status in a short time frame. If I am right, then a smaller ship as the FR could be finding the problems with operating an extreme performance methane engine launch vehicle at a quarter of the engine/airframe cost per boom/lesson. And those lessons could have been accumulating starting about a year ago. This is not casting shade on SpaceX, this is about biting off a huge chunk that may be too much to swallow easily for anybody.

It seems that the strongest argument has to do with Starlink2. If FR were operational by now, it could be launching 100 or so Starlink1 per flight or ~50 of the Starlink2 as they come on line. Assuming that Superheavy/Starship could operate for the same cost per launch as Falcon9, then this purported FR should cost about 25% of that. So Starlink1 would be hitting orbit at roughly 12-13% of the launch cost per sat as those launched on Falcon9. Starlink2 at a little over twice the mass would cost perhaps 30% of what it costs to do a Starlink1 with Falcon9. If Superheavy/Starship has a suggested turn around of up to 3 times per day, then it seems that an FR could hit once per day even while ironing out operational bugs. One ship flying daily with 40 Starlink2 on board would be placing 200 of them per 5 day week. That would be 10,000 Starlink2 per airframe per year. With the current build rate at Boca, there would be no need to rely on one airframe. Perhaps three each at Boca, the Cape, and Vandenberg, with Wallops and Kodiak if needed.

Posted in Uncategorized | 18 Comments

Fill ‘er Up: New AIAA Aerospace America Article on Propellant Depots

Posted on September 2, 2022 by Jonathan Goff

I was interviewed by John Kelvey for an article on propellant depots, which AIAA just published in this month’s Aerospace America magazine. Here’s an online version of the article for those of you who don’t get the print edition. Also interviewed were my friends Laura Forczyk from Astralytical, Jeremy Schiel from Orbitfab, and Bill Notardonato from Eta Space.

Posted in Propellant Depots, Space Policy, Space Transportation | Leave a comment

Independent Perspectives on Cislunar Depotization

Posted on August 26, 2022 by Jonathan Goff

I was invited to participate in a workshop this week in Chantilly, VA focused on cislunar propellant depots. As part of the preparation for the workshop, they asked for stakeholder to provide their perspectives on depot architectures, the benefits of depots, and the technology, market, and policy changes needed to enable them. Based on the inputs I provided, they asked me to share my thoughts with the group.

Over on my Starbright Engineering blog, I wrote a summary of key ideas I think I brought to the table, and key lessons learned:

https://starbrighteng.com/my-independent-perspectives-on-cislunar-depots/

Posted in Lunar Exploration and Development, Propellant Depots, Space Development, Space Transportation | Leave a comment

Starbright Response to ISAM National Strategy RFC

Posted on July 2, 2022 by Jonathan Goff

As I mentioned in my previous post, I’m going to be providing links here to relevant blog posts on my Starbright Engineering LLC blog. This first link is to a comment I submitted to the Office of Science and Technology Policy regarding their In-space Servicing, Assembly, and Manufacturing National Strategy that they released back in April.

Posted in Satellite Servicing, Space Policy, Starbright | Leave a comment

Personal Update: Starbright Engineering LLC

Posted on June 24, 2022 by Jonathan Goff

So, last month I did a thing, and left Voyager to strike out on my own. If you follow me on Twitter, or LinkedIn, or Facebook, you probably already saw. For the near-term, I’ve founded Starbright Engineering LLC as a single-person part-time aerospace/startup consulting business. I’m using Starbright to help pay the bills while I explore other options for my next startup.

I have one idea that I’m exploring with a friend, which is a swing-for-the-fences concept that would take me away from aerospace for a while. It’s highly dependent on negotiating a licensing deal with one or two companies though, so I’d prefer not to go into details publicly yet. My goal is to get to a go/no-go decision on launching this new startup by the end of the summer. If we can negotiate a workable license, and the technology is as near-term as I think it may be, and if we can line up adequate funding to get started, this startup will definitely be a worthy place to invest the next decade of my career. And if we’re successful enough, it may loop back around to aerospace toward the end of that time period.

In the meantime, as I mentioned, I’ll be doing part-time consulting for aerospace and deeptech clients. If you’re interested in talking with me, drop me a line via my Starbright contacts page. I have a non-compete agreement with Voyager, and frankly wouldn’t want to directly compete with my friends companies anyway, so we’ll need to make sure that whatever I’d be doing doesn’t directly compete with anything the Voyager family is trying to do, but it’s often worth discussing to see if there’s a way to angle things to be complementary instead of competitive.

Also, I’m going to try an experiment during this exploratory phase — I’m hoping to be a little more active in blogging in the coming days, but I’m planning on posting blog posts that are professionally oriented on my Starbright Blog, and just posting a link over here on Selenian Boondocks. In the long-run if I decide to shut down Starbright, or if I decide that having yet another blog is a horrible idea, I’ll migrate everything back here. But for now we’ll leave Selenian Boondocks for more speculative posts, and leave more professional/policy-oriented posts for the Starbright blog.

Posted in Administrivia, Starbright | Tagged , , , | 5 Comments

Projectile Fusion

Posted on May 19, 2022 by johnhare

I ran across an article on a variation on impact fusion called projectile fusion. A relatively large (1cm) projectile is smashed into a target at 14,500 mph. The specific shape of the projectile and target temporarily create extreme pressures and temperatures that has created some laboratory detectable fusion reaction. Railgun type accelerators and other complex gear in the UK experiments.

How long before someone thinks to try these experiments in orbit? Retrograde projectiles could easily have a closure rate of 36,000 mph with a prograde target. far beyond the velocities they currently have available on the ground. And far more massive projectiles. Assuming it works, Fusion Orion/Medusa for deep space propulsion without a lot of radioactive mass on board?

Assuming it works, how long before it gets weaponized? Or configured to move asteroids with a large number of small yield devices of even smaller mass?

Posted in Uncategorized | 1 Comment

Payload fraction derivation for vehicle with split delta-V (case #2)

Posted on April 22, 2022 by Kirk Sorensen

Consider a vehicle carrying a payload that undertakes a first $\Delta v$, then drops off that payload and undertakes a second $\Delta v$ in the same overall vehicle configuration (tanks, engines, payload handling, etc.). It carries the propellant for both maneuvers, but only on the first maneuver does it have the added mass of the payload. This situation might be representative of:

1. a reusable lunar lander, based in lunar orbit, fully fueled by propellant delivered from the Earth and loaded with a payload from Earth, which then lands and unloads its payload, then returns to lunar orbit with nearly all its propellant expended.

2. a space tug that departs for geosynchronous orbit carrying a satellite, then returning to a low-Earth orbit for refueling and reloading.

Consider the mass of propellant ($m_\text{prop1}$) for the first $\Delta v$ and the mass of the propellant ($m_\text{prop2}$) for the second $\Delta v$ to be distinct amounts, carried in common tankage.

First, define the mass conditions at the beginning and end of $\Delta v_1$:

(1)   \begin{equation*} \eta_1 \equiv \exp(\Delta v_1/v_e) = \frac{m_\text{vehicle} + m_\text{prop1} + m_\text{prop2} + m_\text{payload}}{m_\text{vehicle} + m_\text{prop2} + m_\text{payload}} \end{equation*}

Alternatively, and just as importantly, the conditions bracketing $\Delta v_1$ can be described in terms of an initial mass:

    \begin{displaymath} \eta_1 = \frac{m_\text{initial}}{m_\text{initial} - m_\text{prop1}} \end{displaymath}

This expression can be conveniently rearranged to yield the propellant mass consumed by the vehicle in $\Delta v_1$:

    \begin{displaymath} m_\text{prop1} = m_\text{initial} \left(1 - \dfrac{1}{\eta_1}\right) \end{displaymath}

In a similar manner, we define the mass conditions at the beginning and end of $\Delta v_2$:

(2)   \begin{equation*} \eta_2 \equiv \exp(\Delta v_2/v_e) = \frac{m_\text{vehicle} + m_\text{prop2}}{m_\text{vehicle}} \end{equation*}

We can also express the conditions bracketing $\Delta v_2$ in another way, in terms of initial mass:

    \begin{displaymath} \eta_2 = \frac{m_\text{initial}/\eta_1 - m_\text{payload}}{m_\text{vehicle}} \end{displaymath}

    \begin{displaymath} \eta_2 m_\text{vehicle} = m_\text{initial}/\eta_1 - m_\text{payload} \end{displaymath}

    \begin{displaymath} \eta_2 m_\text{vehicle} + m_\text{payload} = m_\text{initial}/\eta_1 \end{displaymath}

    \begin{displaymath} m_\text{initial} = \eta_1\eta_2 m_\text{vehicle} + \eta_1 m_\text{payload} \end{displaymath}

The propellant mass consumed by the vehicle in $\Delta v_2$ can also be expressed in a manner analogous to $m_\text{prop1}$:

    \begin{displaymath} m_\text{prop2} = \left(1 - \dfrac{1}{\eta_2}\right)\left(\dfrac{m_\text{initial}}{\eta_1} - m_\text{payload}\right) \end{displaymath}

    \begin{displaymath} m_\text{prop2} = m_\text{initial}\left(\dfrac{1}{\eta_1} - \dfrac{1}{\eta_1\eta_2}\right) - m_\text{payload}\left(1 - \dfrac{1}{\eta_2}\right) \end{displaymath}

Now we are positioned to calculate the total propellant load:

    \begin{displaymath} m_\text{prop} = m_\text{prop1} + m_\text{prop2} \end{displaymath}

substituting the definitions for $m_\text{prop1}$ and $m_\text{prop2}$

    \begin{displaymath} m_\text{prop} = m_\text{initial} \left(1 - \dfrac{1}{\eta_1}\right) + m_\text{initial}\left(\dfrac{1}{\eta_1} - \dfrac{1}{\eta_1\eta_2}\right) - m_\text{payload}\left(1 - \dfrac{1}{\eta_2}\right) \end{displaymath}

collecting terms and simplifying

    \begin{displaymath} m_\text{prop} = m_\text{initial} \left(1 - \dfrac{1}{\eta_1} + \dfrac{1}{\eta_1} - \dfrac{1}{\eta_1\eta_2}\right) - m_\text{payload}\left(1 - \dfrac{1}{\eta_2}\right) \end{displaymath}

    \begin{displaymath} m_\text{prop} = m_\text{initial} \left(1 - \dfrac{1}{\eta_1\eta_2}\right) - m_\text{payload}\left(1 - \dfrac{1}{\eta_2}\right) \end{displaymath}

Now let us define the vehicle’s “dry” mass entirely in terms of initial-mass-sensitive ($\phi$), propellant-mass-sensitive ($\lambda$), and payload-mass-sensitive ($\epsilon$) mass terms. This is a substantial simplification, but it should do for now.

    \begin{displaymath} m_\text{vehicle} = \phi m_\text{initial} + \lambda m_\text{prop} + \epsilon m_\text{payload} \end{displaymath}

    \begin{displaymath} m_\text{initial} = \eta_1\eta_2 m_\text{vehicle} + \eta_1 m_\text{payload} \end{displaymath}

substituting the definition of the vehicle’s mass in

    \begin{displaymath} m_\text{initial} = \eta_1 m_\text{payload} + \eta_1\eta_2(\phi m_\text{initial} + \lambda m_\text{prop} + \epsilon m_\text{payload}) \end{displaymath}

we collect terms related to the initial mass on the left hand side

    \begin{displaymath} m_\text{initial}(1 - \phi\eta_1\eta_2) = m_\text{payload}(\eta_1 + \epsilon\eta_1\eta_2) + \lambda\eta_1\eta_2 m_\text{prop} \end{displaymath}

    \begin{displaymath} \lambda\eta_1\eta_2 m_\text{prop} = \lambda\eta_1\eta_2 m_\text{initial} \left(1 - \dfrac{1}{\eta_1\eta_2}\right) - \lambda\eta_1\eta_2 m_\text{payload}\left(1 - \dfrac{1}{\eta_2}\right) \end{displaymath}

    \begin{displaymath} \lambda\eta_1\eta_2 m_\text{prop} = m_\text{initial} \left(\lambda\eta_1\eta_2 - \dfrac{\lambda\eta_1\eta_2}{\eta_1\eta_2}\right) - \lambda\eta_1\eta_2m_\text{payload}\left(\dfrac{\eta_2 - 1}{\eta_2}\right) \end{displaymath}

    \begin{displaymath} \lambda\eta_1\eta_2 m_\text{prop} = m_\text{initial} \left(\lambda\eta_1\eta_2 - \lambda\right) - m_\text{payload}\lambda\eta_1(\eta_2 - 1) \end{displaymath}

now substituting and collecting terms

    \begin{displaymath} m_\text{initial}(1 - \phi\eta_1\eta_2 - \lambda\eta_1\eta_2 + \lambda) = m_\text{payload}(\eta_1 + \epsilon\eta_1\eta_2 - \lambda\eta_1(\eta_2 - 1)) \end{displaymath}

further simplifying

    \begin{displaymath} m_\text{initial}(1 - (\phi + \lambda)\eta_1\eta_2 + \lambda) = m_\text{payload}(\eta_1(1 + \epsilon\eta_2 - \lambda(\eta_2 - 1))) \end{displaymath}

With all terms relating only to initial mass and payload mass, a general expression for payload fraction can at last be defined:

    \begin{displaymath} \dfrac{m_\text{payload}}{m_\text{initial}} = \dfrac{1 - (\phi + \lambda)\eta_1\eta_2 + \lambda}{\eta_1(1 + \epsilon\eta_2 - \lambda(\eta_2 - 1))} \end{displaymath}

We can compare this to our previous expression for payload fraction by assuming that $\eta_2$ = 1 and simplifying the result.

    \begin{displaymath} \dfrac{m_\text{payload}}{m_\text{initial}} = \dfrac{1 - (\phi + \lambda)\eta_1 + \lambda}{\eta_1(1 + \epsilon)} = \dfrac{\dfrac{1}{\eta_1} - \left(1 - \dfrac{1}{\eta_1}\right)\lambda - \phi}{1 + \epsilon} \end{displaymath}

and see that for the same assumptions they are identical. A bit more insight can be obtained by remembering that the final mass fraction (FMF) is simply the inverse of the mass ratio, and that the propellant mass fraction (PMF) is one minus the final mass fraction:

    \begin{displaymath} FMF \equiv \frac{1}{\eta} \end{displaymath}

    \begin{displaymath} PMF \equiv 1 - FMF = 1 - \frac{1}{\eta} \end{displaymath}

    \begin{displaymath} \dfrac{m_\text{payload}}{m_\text{initial}} = \dfrac{FMF - (PMF)\lambda - \phi}{1 + \epsilon} \end{displaymath}

Remember that this is just for the case where $\Delta v_2 = 0$ and thus $\eta_2 = 1$.

Posted in Rocket Design Theory | 1 Comment

GEO Orbital Debris Mitigation Paper Excerpts

Posted on April 7, 2022 by Kirk Sorensen

Back in 2006, I helped a high-school student (Daniel Rodrigues) who was interested in momentum-exchange tethers to write a paper for a high-school class about a concept for a tether that would remove spent geosynchronous satellites from their orbits quickly, putting them into an elliptical orbit with a perigee that would intersect the atmosphere. He recently recovered the paper and sent it to me at my request, and I am publishing some of the more relevant sections of it here, with minor edits and occasional expansions for explanation.

Background on Momentum Exchange Tethers

The Momentum Exchange Tether is a concept originally pioneered by Hans Moravec in 1977[1]. This orbital facility, essentially a rotating cable in orbit of the planet, had the ability to touch the surface of the planet every 20 minutes, and lift payloads into orbit. Carroll, in 1991, evolved this design into a totally in-space system able to lift payloads from sub-orbital trajectories and toss them into higher orbits [2]. The Momentum Exchange Tether system was refined further in 1998 by Bangham, Lorenzini, and Vestal, who designed the system to transfer payloads from LEO to Geostationary Transfer Orbit (GTO), and who concluded that the system should be composed of two separate facilities: one at an altitude of 2019 kilometers, and another at 25048 kilometers [3]. In 1999, Hoyt and Uphoff reestablished the one tether design due to simplicity concerns, but retained the LEO to GTO configuration[4]. Since this study, the Momentum Exchange tether has been refined by Hoyt once again in 2000[5], by Sorensen et. al. in 2003[6], and finally Hoyt once more, also in 2003[7]. As of now, the standard Momentum exchange tether is situated in a GTO, rotating so that its angular velocity and orbital velocity, when combined, equal the orbital velocity of it’s payload in LEO. At the tether’s perigee, it rendezvouses with the payload, rotates 180 degrees, and releases the payload. This adds momentum to the payload, but takes it away from the facility, consequently lowering its orbit. However, ballast at one end of the tether disallows for a significant drop in altitude. The station is then reboosted, and is then ready for another payload. The largest portion of the system, the cabling, is composed of a series of interlocking primary and secondary lines, a design known as the Hoytether [19].

The Problem with Orbital Debris

This investigation aims to apply the momentum exchange concept to the deorbiting of unwanted satellites (otherwise known as orbital debris, or simply debris). Orbital debris consists of inactive spacecraft, spent rocket stages, spacecraft fragments, and other miscellaneous objects [8]. Objects in the .01 to 1 cm size category can cause significant system damage, and objects larger than a centimeter can conceivably be catastrophic [8]. Additionally, spacecraft can only be shielded against debris up to 1 cm, due to mass practicalities [8]. Therefore, it can be concluded that orbital debris is a considerable threat to spacecraft.

Orbital debris can be especially troublesome in Geosynchronous Earth Orbit (or GEO), a very valuable region in space. At GEO altitude (35,678 km above the Earth’s surface) satellites orbit the Earth at the same rate as the Earth rotates. This means that the satellites stay over a fixed point on the planet, which as obvious commercial value for communication satellites. However, this region in space is rapidly filling with debris. As more satellites are put into GEO, the amount of debris only increases. As of January, 2005 there were 153 tracked, uncontrolled objects in GEO, or about 14% of all known objects in GEO [9]. These uncontrolled objects will only endanger the, on average, 21.1 satellites that will be launched annually into GEO from now into the near future, which is a number certain to increase in time [10].

Objective Statement

A momentum exchange tether will be designed that is capable of capturing uncontrolled satellites and orbital debris of moderate dimensions in Geostationary Earth Orbit (GEO) with a maximum mass of 5400 kilograms (encompassing 83% of satellites expected to be launched, into 2013)[10], and propelling the debris to a negative perigee altitude, ensuring the destruction of the debris. The momentum exchange tether will also be capable of capturing satellites in Geostationary Transfer Orbit (GTO), and slinging the payload into Geostationary Earth Orbit, losing momentum gained as a consequence of deorbiting debris.

Initial Calculations

The design of the tether system began with the analysis of the change in velocity required to remove a satellite from GEO into a deorbit ellipse intersecting with Earth’s atmosphere before perigee. The perigee of this ellipse was selected to be at an altitude of -100 km. This was done for several reasons: First, it ensured the destruction of the debris with a moderate margin for error, and second, it allowed for some flexibility in the coming design stages. The delta-v was set up between the initial orbit of the debris, GEO (circular equatorial orbit with an altitude of 35786 kilometers) and the target ellipse, which intersected with the Earth’s atmoshere. An infinitely massive ballast was also assumed in these initial, preliminary calculations, in order to simplify the center of mass variables. To calculate the speed of a satellite in GEO, the following equation was used:

This equation solves for the tangential velocity of an orbiting body (v), taking into account the Gravitational Parameter (mu) and the radius of the orbit (r). Afterward, the specific mechanical energy of the target ellipse was calculated. The specific mechanical energy (epsilon) is essentially a value that gives the energy per unit of mass of an orbiting body. The equation for this value is as follows:

where “a” is the semi-major axis. From the specific mechanical energy, tangential velocity of an orbiting body can be calculated using an alternate from of the equation for specific mechanical energy:

This equation was used to calculate the tangential velocity of the payload at the apogee of the target ellipse (after release from the tether). When the velocity of the payload in GEO is subtracted from this value, a new value of -1.5017 km/sec is calculated. This means that a vector with a magnitude of 1.5017 km/sec must be applied to the payload, in the opposite direction of the path of the payload, which accounts for the negative value.

Tip Velocity and Rendezvous

Once it was calculated how much the debris needed to be slowed down by (1.5017 km/sec), the actual tether facility could be designed. The first step in this process was to determine how the tether would perform its intended task. In this particular configuration, the tether would be situated below its payload. The tether’s center of mass (CM), about which it rotated, would be traveling in the same direction as the payload, albeit slower. The tip of the tether configured to capture the debris would also rotate in the direction the debris was traveling. Capture would occur with the payload directly above the tether’s CM, when the tether reached its apogee. The tether would then release the debris one-half of a rotation later, decelerating the debris into the above-discussed deorbit ellipse. Possibly the most critical value in this scenario would be the tether’s tip velocity, which is the velocity the tether would impose onto its payload. As mentioned previously, the debris needed to be slowed by 1.5017 km/sec. But because the tether imparts this velocity onto the payload in two ways (this will be elaborated upon in a moment), this value was divided by two, which gave the value of 0.75085 km/sec as a tip velocity.

Rendezvous between the payload and tether tip would not be possible if the velocities were not matched [6]. Therefore to find the speed of the tether’s CM, the tip velocity was subtracted from the velocity of the debris (3.074 km/sec). Therefore, the combination of the tether’s CM vector and the tip vector would equal the vector of the debris. In this scenario, the tether’s CM was calculated to be traveling at 2.32315 km/sec, at apogee. From the debris’ point of view, because vectors along the x-axis would be matched, the tip of the tether would be approaching from below it, along the y-axis. When the payload is released, 180 degrees later, the tip vector is once again added to the tether’s CM vector. However the tip vector is not 0.75085 km/sec, but is now -0.75085 km/sec. Therefore, the payload is released at the desired velocity of 1.5017 km/sec, and enters the deorbit ellipse. Also, it is now clear why the tip velocity was equal to the desired velocity divided by two: the other half of the velocity to be imposed on the payload came from the CM vector. This allowed for a capture with a 0 km/sec relative velocity when the tip was rotating with the debris and the CM, and allowed for the desired velocity of 1.5017 to be attained once the CM and tip were traveling in opposite directions. It is important to note that because an infinitely massive ballast mass was assumed in these initial calculations, actual velocities before capture would be different. However, these variables would change commensurately, retaining the 0 km/sec relative velocity.

Ballast

The next component that must be analyzed is the ballast mass. The ballast serves as a mass that stores momentum, allowing for smaller changes in altitude after release [7]. The simplest and most efficient way to fix a ballast mass to a tether station is to utilize the rocket the station was launched in, which lowers initial launch cost by preventing the launch of additional mass to be used as ballast[7]. In this study, the Ariane 5-ECA launch vehicle was chosen as the most suitable rocket. Its high capacity (10,500 kg[14]) allows for massive satellites such as the tether station to be launched, and it has the capability to launch into a GTO [15]. The upper stage on the vehicle, the ESC-A, has a dry mass of 4540 kg [15], which is therefore the mass of the ballast.

Capture Mechanism

On the opposite side of the station is the capture device. The purpose of this device is to physically anchor the payload (in this case, orbital debris) to the tether itself, once position and velocity is matched. The design of the device must incorporate some margin of error. A mass of 200kg was estimated for this device, which will be capable of securing payloads of at least 1 meter long in any dimension. This is because debris in GEO that is tracked must be at least a meter wide to be traced, due to limitations in radar technology [10].

Concept of Operations

The facility will be launched in an Ariane 5-ECA rocket equipped with an ESC-A upper stage, into its orbit. The tether can then be fully deployed, and spun up to an angular velocity of 0.01882 radians/sec. This value will remain constant. The system is now ready for operation. The angular velocity of the system gives the tip a tangential velocity of 1.0921 km/sec. Added to the CM velocity of 1.9819 km/sec characteristic to the system’s current orbit will yield a value of 3.074 km/sec, identical to the orbital velocity of debris in GEO. After capture of debris, the CM of the tether shifts, toward the tip. This slows the tip velocity to 0.75085 km/sec, while accelerating the CM velocity to 2.32315 km/sec. This is because the tangential velocity of the point where the CM was shifting to is added to the old CM orbital velocity. This acceleration causes the altitude of the facility to increase, and is the first illustration of momentum exchange. After half a rotation, the debris is released. Once again, the CM shifts up, and accelerates, increasing its velocity. This also increases the station’s altitude, just as the previous maneuver. After release, the debris enters its deorbit ellipse. The velocity of the debris is now 1.5723 km/sec, at a radial distance (from the Earth’s center) of 42066.2 km. The debris under these circumstances would travel in an ellipse with a perigee altitude of -67.5 km. This means that the debris would enter the Earth’s atmosphere before its minimum altitude, allowing for a significant margin of error.

After release, the tether would be in a significantly higher state of energy, and therefore a higher orbit. The tether must be brought back to an altitude where it can perform its task of deorbiting unwanted space junk. There is a nearly “free,” and almost instantaneous way to bring down the facility. This would involve capturing a functional satellite on its way to GEO in a GTO intersecting with the fully boosted tether facility. The new satellite would interact with the tether in the exactly opposite way as the debris, as long as it has the same mass as the debris that was deorbited. If not, on board thrusters could complete the slowing of the facility to its original orbit. But inserting a functional satellite into GEO would save a significant amount of propellant, and would eliminate the need for an entire upper stage on the new satellite, saving more money. On a 5400 kg satellite, almost 2000 kg is used as on-board propellant and thrusters. This would mean a reduction in launch costs from Earth into GTO, which is typically about $10,000 per pound[16].

Of course, sending debris crashing through Earth’s atmosphere raises the question of safety. However, reentry of space debris is a very common occurrence. In the past 40 years, there have been over 16,000 known re-entries of cataloged space objects, without significant damage or injury[17]. This is due to both the fact that most if not all of the debris disintegrates in the atmosphere, and the fact that any remnants of the doomed spacecraft have a very low probability of impacting populated areas[18]. Regardless of the unlikeliness of a ground impact in a populated area, the decision to deorbit any particular non-functional spacecraft will have to be made on a case-by-case basis. Large objects have been known to survive reentry in past, and the reentry of any spacecraft containing radioactive substances is out of the question[18].

Conclusions

This investigation has outlined the basic configuration of a momentum exchange tether in GTO capable of both deorbiting debris, and putting new satellites in GEO. This tether will increase lifetimes of satellites in GEO by reducing the threat of debris, while reducing the cost of launching new satellites. Orbital debris is accumulating rapidly, and a solution such as the momentum exchange tether needs to be considered. GEO is a valuable natural resource that needs to be conserved, just as any other. The combination of a less hazardous environment in space with lower launch costs is certain to stimulate the development of a stronger space infrastructure, undeniably helping humanity expand its horizons.

References

1.Hans Moravec, A Non-Synchronous Orbital Skyhook, AI Lab, Computer Science Dept., Stanford University, Stanford, Ca. 94305

2.Carroll, Preliminary Design of a 1 km/sec Tether Transport Facility, March 1991, Tether Applications Final Report on NASA Contract NASW-4461 with NASA/HQ
3.Bangham, Lorenzini, Vestal, Tether Transportation System Study, NASA TP-1998-206959, 1998

4.Hoyt, Uphoff, Cislunar Tether Transportation System. AIAA 99-2690, 1999

5.Hoyt, Design and Simulation of a Tether Boost Facility for LEO-GTO Transport, Tethers Unlimited, Inc., Seattle, WA, AIAA 2000-3866, 2000

6.Sorensen et. al., Momentum eXchange Electrodynamic Reboost (MXER) Tether Technology Assessment Group Final Report, NASA Marshall Space Flight Center, 2003

7.Hoyt, Slostad, Frank, A Modular Momentum-Exchange/Electrodynamic-Reboost Tether System Architecture, AIAA-2003-5214, 2003

8.Interagency Report on Orbital Debris, The National Science and Technology Council, Committee on Transportation Research and Development, 1995

9.Serraller, Classification of Geosynchronous Objects Issue 7, European Space Agency, January 2005

10.2004 Commercial Space Transportation Forecasts, Federal Aviation Administration’s Associate Administrator for Commercial Space Transportation and the Commercial Space Transportation Advisory Committee, May 2004

11.Pro Fiber Zylon, Toyobo Co., LTD., 2001

12.Personal Correspondence with Kirk Sorensen, In-Space Propulsion Technology Projects Office, NASA Marshall Space Flight Center, June 22, 2005

13.Sorensen, Conceptual Design and Analysis of an MXER Tether Boost Station, Propulsion Research Center, NASA Marshall Space Flight Center, AL, AIAA 2001-3915, 2001

14.Technical Information Ariane 5, Arianespace International Affairs and Corporate Communications, Arianespace, 1999

15. Ariane 5 Users Manual Issue 4 Revision 0, Arianespace, Courcouronnes, France, November 2004

16. Futron Corp., Space Transportation Costs: Trends in Price Per Pound to Orbit 1990-2000, Sep. 6th, 2002

17.Technical Report on Space Debris, United Nations Scientific and Technical Subcommittee (STSC), New York, ISBN 92-1-100813-1, 1999

18. R. P. Patera, and W. H. Ailor. The Realities of Reentry Disposal, A98-43901 12-12, Spaceflight Mechanics 1998: Proceedings of the AAS/AIAA Space Flight Mechanics Meeting, Monterey, California, February 9–11, 1998

19.Foward, Hoyt, Failsafe Multiline Hoytether Lifetimes, 31st Joint Propulsion Conference and Exhibit, AIAA 95-2890, 1995

20.Orbital Debris: A Technical Assessment, National Academy Press, 1995

21.Grun et. al., Collisional Balance of the Meteoritic Complex, Icarus, 1985

22.Cour-Palais, B. G., Meteoroid Environment Model-1969, NASA SP-8013, 1969
​

Posted in Orbital Debris Remediation/Mitigation, Orbital Dynamics, Space Tethers | 1 Comment