Starbright Response to ISAM National Strategy RFC

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

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

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)

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

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

L4 Opportunity

https://www.space.com/earth-extra-moon-trojan-asteroid-2020-xl5-discovery

An Asteroid 1.2 km across has been discovered at the Earth Sun L4 region. This is the second asteroid discovered in one of the Lagrange volumes of the Earth Sun with the other one about a third of that size. It is suggested that it is possible that there are more, quantity unknown, bodies in those regions. They are very hard to locate from Earth due to the distance and sun angle so actual number, makeup, and size in each L region is mostly speculation.

I think it might be possible that this is an ideal target for in situ resource utilization. The advances in cubesat capabilities open many new possibilities. A lander could test out a number of resource extraction techniques for immediate use. Regolith shielding? Steam rocket with whatever volatiles are located? Sintered structure in vacuum and low gravity? Low tech metal extrusion for structural components? Low tech film Solar sails?

The purpose would be to use it as a base to locate other materials in the region with a relatively small scope. And go visit them if found using local resources. Prospecting and exploring and evaluating. If there does turn out to be a large number of bodies in the L4 region, DeltaV per visit should be very low. This one is a C type, dark and carbon. There might just be enough variation in different bodies in the area to make it a resource rich destination.

How simple and cheap could a probe be and still investigate the possibilities??

Posted in Uncategorized | 5 Comments

Random Thoughts: A Joint International Debris Remediation Effort

Galactic trash orbiting Earth
We Don’t Want This to Happen (Credit: Getty)

As humanity’s ambitions in space increase, one of the biggest potential risks to to those ambitions is the dangers caused by the debris we’ve left on orbit over the first half century of space development. Since Sputnik’s launch in October of 1957, we’ve left over 18,000 pieces of debris on orbit that are larger than 10cm, and millions of pieces too small to track but big enough to destroy a satellite. One of the biggest threats comes from large satellites and more especially rocket bodies that were left on orbit during the early days of space development before countries properly realized the risks that leaving debris on orbit could pose. At last year’s IAC conference, a joint group of researchers from 11 countries published a list of the 50 most dangerous pieces of space debris on-orbit, and the vast majority of those 50 were rocket bodies, with most of them being ones launched by the now no-longer existent USSR.

Why Should We Care About Large Debris?

Derelict rocket bodies and large satellites have been known to breakup on-orbit, sometimes due to pressure vessels or batteries failing, and those explosions can create hundreds or thousands of new pieces of debris1. More importantly, derelict space objects can’t dodge, so while we can track them, there’s not much we can currently do if two school bus-sized derelict space objects are on a collision path at relative velocities several times faster than the fastest bullet2.

As Joe Carrol once put it, the best way to avoid creating “BBs” (untrackable but lethal space debris) and “hubcaps” (barely trackable lethal debris) is to get rid of the derelict school buses. If we want to see a world with multiple commercial LEO facilities, propellant depots, space settlements, and especially things like large fleets of Starship-sized spacecraft heading out to Mars or other destinations, we can’t allow the Low-Earth Orbit environment to become a shooting gallery of space debris.

What’s Standing in the Way of Solving this Problem?

Almost all of the most dangerous space debris up there was launched by governments as part of civil or military space missions. As such, it makes sense that governments should pay to clean up the environmental mess their activities created. While there are now early efforts in Europe and Japan to begin tackling Active Debris Removal of dangerous space debris launched by Europe and Japan, that still leaves most of the most dangerous pieces of space debris currently unaddressed.

As I understand it from conversations with my friends in the Space Law community3, a big part of the problem with addressing many of the most dangerous pieces of space debris is that as part of the Outer Space Treaty, which governs the actions of all major spacefaring nations, you’re not allowed to interfere with objects launched by other countries without their permission. Unlike on the oceans, there is currently no space equivalent of salvage law, or “flotsam and jetsam”. Once an object is launched by an actor within a state, that state retains responsible for those objects indefinitely. However, as I understand it, those states are only liable for damages if the object deorbits and damages something or hurts someone on the ground — if it damages something in space, holding them accountable would be really hard unless you could prove that they violated the accepted standards of care at the time they last had control of the object. So, as I understand it, most of the most dangerous space debris is owned by a country that can’t realistically pay to clean it up, can’t be deorbited by someone else without the original country’s express permission4, and space liability law won’t actually hold the owner responsible if that debris creates a ton of new debris, so they don’t have any incentive to clean things up, other than wanting to be a good citizen.

Additionally, the US government has been very reticent to fund developments in Active Debris Removal, because of the fear that adversaries might see these technologies as dual-use technology5, causing them to invest in similar technologies that could be used against the US.

Anyhow, at a World Economic Forum meeting on the topic of space debris and space traffic management, someone floated an idea for a potential way to solve these problems. I can’t remember who suggested the idea, to give proper credit, but I wanted to run with it, and try to bake it a little more6. The idea is what if we propose to clean up the most dangerous pieces of large debris via a joint international cleanup effort?

Concept for a Joint International Debris Remediation Effort

While we’re on the cusp of having the technology to solve the problem of large space object active debris removal (ADR), the policy and international relations issues are the bigger unsolved pieces of the puzzle. So the idea is, what if we find a way to jointly tackle the problem in a way that incentivizes all of the key players to act, and removes some of their bigger concerns about acting?

What I’m Proposing: A joint effort, involving at least the US and Russia, but possibly also China, Japan, and Europe, to capture and recycle7, and/or controllably dispose of the most dangerous 100-200 pieces of space debris.

Key Elements of the Concept:

  • Most of the funding for the effort would need to come from the wealthier spacefaring nations. We have the wealth to do this, and stand the most to lose from inaction.
  • To make it worth Russia’s while to participate, there should be a way for them to economically benefit from this effort — my suggestion would be to have them involved at least with the on-orbit recycling part of the effort. Helping them build up new and useful space economic capabilities and industries is a good incentive for cooperation. Though in theory, the funding states could maybe insist on these being joint ventures between companies in their countries and Russian companies, as a way to share in the upside.
  • If China also wants to participate in this a mutually-acceptable portion of the project could be done via Chinese companies, or joint ventures with Chinese companies.
  • It’s a good idea to have dissimilarly redundant RPO sensors, especially for something as dangerous as capturing and detumbling a spinning rocket stage. So, it might be possible to have both commercial-grade US and Russian (and Chinese?) RPO sensors on the vehicle, with the feeds from both of them being live streamed during RPO events for enhanced transparency8
  • Having the most sensitive operations, like the RPO and detumbling/grappling efforts, lead by commercial companies in more neutral countries might also help. Companies from the US, Russia, or others could still develop technologies for use on these missions, but have them operated by someone that is as innocuous as possible to both US and Russia. Conveniently, two of the main companies working on active debris remediation are a Swiss company, Clearspace9, and the Singaporean/Japanese/US/UK/Israeli company Astroscale.

There are probably plenty of other elements here, but I wanted to get the idea out there. A joint US/Russia or US/Russia/China mission like this could be a way to break the logjam on cleaning up the space environment before the situation degrades further, while also helping greatly accelerate development in advanced on-orbit operations that could help enable not only in-space recycling, but also in-space construction and eventually asteroid mining.

What do you all think?

Posted in Business, Commercial Space, International Space Collaboration, International Space Competition, Orbital Debris Remediation/Mitigation, Random Thoughts, Satellite Servicing | 3 Comments

Suborbital Balloot

Watched a bit of the two suborbital flights, not that much though, certainly not with the interest I would have had roughly a decade and a half ago. I had thought back then that the technical route to routine space travel would be through incremental advancements from suborbital to orbital to cislunar to other planets and asteroids. Looks like I was wrong again.

One thing that stood out to me was the people trying to play in microgravity with limited room in which to do it. Much more limited in volume and more clutter than on the ZeroGee flights. Looked to me like time was limited enough that there was just time to glance out the windows and play half a round of ZeeGee catch.

A balloot in the spirit of the inflatable space structures might help here if some problems of fast deployment, rapid access, and safety could be addressed. A 5-10 meter diameter inflatable without seats and other clutter would give considerably more play room than either SS2 or NS. Fast deployment might be a bit of a challenge as there wouldn’t be time for a leisurely five minute deployment and check out.

I don’t see how the idea could be adapted to SS2, but it does seem that a VTVL ship could have a fast clamshell hatch in the nose that could allow people in a fairly tight cabin access to a large play area within seconds of reaching vacuum and engine cut off. So a ship that would have had six people might fly with a dozen, each of which would have a lot more room to play and experiment.

With a large balloot starting deceleration at much higher altitude, it seems possible that a ship could fly much higher without subjecting the participants to excessive gee loads. Every extra Mach number gives on the order of another minute of play time in microgravity. On the way back, the participants get back to their seats while the hatch closes. The balloot remains deployed until landing as a drag device incidentally reducing terminal velocity considerably, possibly enough to make a failure of engine relight survivable, if painful.

This seems to me an idea that, even if feasible, is very late to the party. It looks like orbital tourism is likely to become fully operational in the same timeframe as the two suborbital contenders. The contrast may well cut into the desirability of the short pop ups. One interesting factoid though is that the people that were insisting that orbital was 64 times as hard as suborbital have to be wondering what the two companies could have accomplished with 64 times the investment.

Posted in Uncategorized | 9 Comments

Layover Lunar Lander

A lunar lander concept I haven’t seen elsewhere is a layover concept. A tall ship lands on its’ tail the way the several suborbital hoppers and the Falcon9 has been doing for several years now. After landing and checking the area to one side, a small thruster tips it over on two legs. Before reaching the surface, some landing rockets in the nose fire to bring velocity to zero as the nose lands on a couple more much smaller legs.

If the lander is 100 meters tall, the terminal velocity that must be braked before the nose landing is under 20 m/s. Stopping 100 tons just at the surface would require between 400 and 700 kg of propellant depending on the Isp in use. If feasible, seems like a fair trade off as opposed to raising and lowering people, supplies, and equipment 100 meters for the duration of the mission.

The lower legs on the pivot side would need to be able to handle loads from two directions, though the legs at the nose could be quite light. One alternative leg option is to have the lower pivoting legs articulate up under the body of the vehicle to split the loading better and to reduce braking thrust requirements a bit more.

The tip over landing should be quite simple as there would be no question of where the vehicle is in relation to the ground at any given time. Of course smaller vehicles would have even less trouble and use less propellant. A 20 meter high vehicle with a 10 ton payload would require only 20-30 kg of braking propellant.

Posted in Uncategorized | 24 Comments

Unorthodox Reusable Lunar Landers Concepts

[Author’s Note: Back in summer of 2019, shortly after Mike Pence announced the goal of having NASA return astronauts to the lunar surface for the first time in over half a century, I had the idea of doing a blog post about the benefits of fully-reusable lunar landers, and then going over a few of my favorite unorthodox reusable lunar lander configurations. I got side-tracked at the time by my entrepreneurial day job1, and by the time I freed up from that, Altius was busy supporting one of the three Artemis HLS teams, so I felt a blog post like this might be impolitic. Now that we’re no longer actively supporting that team, and given that I no longer have running a startup as an excuse for not blogging, I wanted to finish fleshing out these ideas and at least get them out there for discussion.]

While traditionally most lunar landers and lander conceptual designs have been fully-expendable, many people, including NASA have begun to see some of the benefits of reusable lunar lander systems. Some of those benefits as I see them include:

  • Lowers the cost of sending hardware to the Moon: This is the obvious one that people get about reusability, is that so long as it’s done right, it can significantly lower the per mission cost, especially if the lander can be reused many times over its lifetime.
  • Makes the program easier to throttle up or down: One of the big challenges the Augustine Commission noticed for NASA Human Spaceflight missions was that they could rarely afford to both develop new capabilities while operating previous ones. With a fully-reusable lander system, especially one designed to not require a huge standing arm to support it, much of the cost of a given mission could be the marginal cost of launching new crew/cargo/propellant, which means it’s easier to throttle down temporarily without losing the capability.
  • More amenable to non-critical-path international participation: With reusable landers, once they’re launched, incremental missions mostly require refueling, reprovisioning, and a crew swapout. Government space agencies typically don’t want to spend money outside of their country–they typically try to find a way where they can handle things through barter and no-exchange-of-funds agreements. For instance, in exchange for getting some of the ISS crew slots, Japan and ESA both developed cargo vehicles to bring some of the cargo that ISS needed. So they could spend their space agency money locally, and use that contribution to get ISS astronaut slots without having to develop their own crew launch capabilities2. Reusable lunar landers provide an easy barter option for international participation in NASA lunar missions — launching propellant and/or cargo. The nice thing is that if done wisely (say using a Low-Orbit HSF Depot in LEO), this might require minimal development cost for foreign countries, while allowing them to usefully contribute, but in a way where they’re not on the critical path, and they can throttle up or down their involvement as desired.
  • Creates near-term demand for lunar ISRU: Once you have reusable lunar landers, the vast majority of the mass needed per mission is propellant. Being able to source that locally could significantly reduce the cost of missions3, and could increase the capability of landers by enabling them to be refueled both in orbit and on the surface4. With reusable landers, you have established demand at an established price point, which makes closing the business case for lunar ISRU easier, so long as you can truly extract it cheaper than shipping it from Earth5.
  • Enables a much more ambitious exploration program: This should be obvious, but once you have reusable landers, you have tons more flexibility for doing things beyond simple flags-and-footprints missions. Things like lunar search and rescue, doing suborbital sorties from a bigger outpost to explore areas of economic or scientific interest, etc. become more feasible.

Anyhow, if you’re reading this blog, I’m probably preaching to the choir here, but I wanted to lay out some of my thoughts on why reusable lunar landers matter.

Unorthodox Reusable Lander Concepts

Given the desirability of fully-reusable lunar landers, it’s sad that most of the best-known reusable lander concepts have used a very similar landing configuration — tall skinny landers with the crew and cargo mounted on top.

Traditional Reusable Lander Concepts: One Long Elevator Ride for a Man, One Griant Crane for Cargo… (credits: SpaceX, Chesley Bonestell x2, Lockheed Martin)

In this series I wanted to highlight a few other potential fully-reusable lander configurations worth considering, some thoughts on variations on the themes, and their pros and cons. Only maybe one of these configurations is one I could claim to have invented, but I thought it would be worth highlighting some other good ideas that may or may not be as well known, especially to younger space engineers or enthusiasts.

In the following parts I’d like to discuss the following reusable lunar lander configurations:

  1. Bottom-loader SSTOs
  2. Horizontal-landing SSTOs
  3. Un-Crasher TSTOs

I may think of and add additional configurations later if time permits. But next up: Bottom-loader SSTO landers.

Posted in ISRU, Lunar Commerce, Lunar Exploration and Development, Propellant Depots, Reusable Lunar Landers | 8 Comments