33

Imagine a large vehicle that can recycle water and air, has shielding, and contains everything for a trip that would take months. This would be the cycler. The vehicle is heavy and is needed for a long trip; crew cannot survive in a small rocket. The benefit is that when you launch your crewed rocket to catch up with the cycler, you only need to accelerate ...


27

One good thing about these periodic trajectories / interplanetary orbits, be it Mars Cyclers, Earth-Moon Cyclers (gravity assist UP/DOWN Escalator Orbits), or Resonant Cyclers (fixed VISIT Orbits) is that they can be maintained with a relatively small penalty for the total mass of the spacecraft, so they could be whole industry complexes processing raw goods ...


17

The question is somewhat odd. The "Interplanetary Transport Network" may be a misleading term. When probes are sent into deep space, most of them make use of flybys or gravity assist manoeuvres. Virtually every celestial body can therefore be used for increasing the speed of a probe or decreasing it. The "network" refers to series of such manoeuvres. In the ...


12

No, a small push will not change a satellite's inclination by 90°. Remember that, when in low Earth orbit, a satellite is already travelling some 7-8km/s. A small push of only some m/s will leave the satellite with almost exactly the same trajectory, changing its inclination by only a small fraction of a degree. In order to change your inclination by a ...


12

Is it possible to create a cycler that can travel from Venus to Mars, then back to Venus? Yes. When only two planets are involved, there are several infinite families of cycler orbits. Given the 434 day long orbital period of an ellipse touching both Venus and Mars is almost twice as long as the 225 days orbital period of Venus, it's even possible to ...


11

How do you minimize Δv to get to Mars? The answer is simple: Wait until 2018 or 2035. Those are the local minima in the Δv needed to get to Mars. The required minimum Δv varies by a large amount. There's a local minimum roughly every two years where transfers to Mars become feasible, but even this quantity varies considerably. There's a ~15 year variation ...


10

I don't think that's really the question you want to ask. The lowest energy trajectories are ones like you suggest, which achieve most of the energy to get to Mars with short impulsive burns very close to Earth. That requires chemical or nuclear thermal propulsion systems that can expend much of the propellant over a very short time. What you want for ...


10

The sentence reads "Earth-Mars cycler applications generally suffer from long repeat periods, infrequent launch opportunities, and the risky requirement to perform hyperbolic rendezvous" The bolded portions are important parts of that sentence. If a rendezvous is botched during planetary fly by, there won't be another opportunity until next fly by. The ...


9

The Lagrange 1 and Lagrange 2 regions are the hubs of the trajectories that Ross, Lo, Martin, and Belbruno advocate. Szebehely developed equations to find the L1 and L2 regions. Using his equations I made a L1 and L2 spreadsheet. Of interest is the 3 body mass parameter, usually denoted $\mu$. $\mu=mass_{orbiting body}/(mass_{orbiting body} + mass_{central ...


8

It is my belief that a bitangential transfer between two coplanar orbits is best. For such a transfer, no direction change is needed at departure or arrival since velocity vectors are parallel at these two points. The red orbit pictured above is tangent to the circular departure orbit as well as the elliptical destination orbit. Here's a picture of many ...


7

A connection between two different periodic orbits around Lagrange / libration points is often termed a heteroclinic connection. Although the L1 and L2 points in the Earth-Moon system are themselves unstable equilibrium points, there are stable periodic (or quasi-periodic) orbits that can be found around each point. From these orbits, it is possible to ...


7

Page 10 of the article: "Earth-return trajectory that has a period of 1/2 month (or 1/3 month is some cases) in order to ensure return to the Moon after 2 (or 3) revolutions of the Cycler in its Earth-return orbit". Those revolutions are not shown in the diagram; "Earth Return Trajectory" and "Translunar Trajectory" only show half of their respective ...


7

Coming from one of the sources, a hyperbolic trajectory rendezvous is dangerous because if the rendezvous doesn't do as planned, the cycler taxi will escape the body's gravitational pull and venture out into space. From the paper: "Apollo missions included a similar risk when the lunar module docked with the command/service module in lunar orbit. If this ...


7

I found this kind of counterintuitive as well at first. This is the way I rationalized it to myself: consider a high-thrust maneuver to change inclination. Obviously if it's impulsive, you perform it at the node. If it lasts for, say, one minute, you would burn +/- 30 seconds around the node. Now, take that to the limit, where the burn time is the entire ...


5

To enjoy a bi-elliptic savings, destination needs to be greater than destination orbit by a factor of 11.94 or greater. Going from LEO to the moon certainly qualifies: 384,400/6678 = ~58. I'll pull a bi-elliptic path out of the hat, don't know if it's optimal: 300 km altitude circular orbit to a 900,000 km altitude apogee: 3.16 km/s Apogee burn at 900,000 ...


5

There is a tool "LTool" by Martin Lo from JPL to compute low-energy transfers via chaotic trajectories near Lagrange Points. Martin W. Lo and Roby S. Wilson The LTool Package http://www.ieec.fcr.es/libpoint/abstracts/lo2.pdf or copy http://www.ieec.cat/hosted/web-libpoint/abstracts/lo2.pdf Martin Lo, LTool Version 1.0G delivery memorandum // JPL TRS 1992+, ...


4

Wow, three years and still no answers!. I will give it a try. What I am going to answer applies just for transfers between periodic orbits at the Lagrangian points (the OP asked so many things but I think that is the fundamental one) Let's put that we want to transfer from Earth-Moon L1 to L2 using the CR3BP dynamics Define the periodic orbits around L1 ...


4

According to this recent Scientific American article, there is a cheaper way to get to Mars than the traditional Hohmann transfer. The mission design relies on a ballistic capture to do away with the Mars orbital insertion burn. The premise of a ballistic capture: Instead of shooting for the location Mars will be in its orbit where the spacecraft will ...


4

Any payload which is supposed to travel with it would first need to accelerate a lot to rendezvous with the cycler and then, at the destination, decelerate a lot to rendezvous with the planet. There is some work that has been done to find optimal cycler orbits, check this work. Some of the computed orbits reduced the delta-V for inbound taxis almost to ...


4

I can give you an upper bound for the necessary $\Delta v$, between all elliptical orbits, regardless of inclination. You can always do a bi-elliptical transfer, performed by almost reaching escape velocity, then do a close to 0 $\Delta v$ manoeuvre at infinity, and then fall back to insert into the target orbit. Both the large burns are done at periapse. ...


3

Set your time and length units to LD and lunar period and you can lose the 2 pi and mu. $T=a^{3/2}$ So if you wanted an ellipse whose period was 2/3 that of the moon you'd have $2/3=a^{3/2}$ $2/3^{2/3}=a$ $a=.763$ .763 of an LD is .763 * 384400 = 293352 km. If perigee were 6678 km then apogee would be (2*293352)-6678. Which is 580026 km. If you wanted ...


3

The formulas (15) and (17) are given under the assumption that both $f_1$ and $f_2$ (the radial and the transverse components of the perturbing force) stay "sensibly constant over one orbital period". In particular, it means that the angle between the force and the radius vector stays sensibly constant. The two cases you are talking about do not satisfy this ...


3

According to Newton, for every action there is an equal and opposite reaction. As the magnets were applied to the craft, it would apply as much force on the accelerator. The accelerator would have to be very massive and/or have rocket engines to generally maintain position and maximize the force imparted on the spacecraft. Of course assembling such an ...


3

This is a well-known phenomenon in helicopter dynamics and in control systems. In a second order control system (which roughly describes your low-thrust scheme), the phase angle change is 90° when the input control frequency is the same as the system's natural frequency. (See any elementary control text and look at the second order system frequency response.)...


3

Usable Mercury cyclers exist. Given the large inclination of Mercury, we want the encounter to happen along the line of apsides, thereby requiring a quasi-periodic stationary cycler. The simplest type of cycler in this family is a high periapse fly by ellipse, ideally close to a Hohmann transfer orbit. This requires that the planets' synodic period is a ...


3

In addition to possibilities outlined in your post, you can also take advantage of the fact that (gross simplification ahead) orbits around Lagrange points are possible, but very unstable, letting you "choose the instability you want" and leave the Lagrange point in almost any direction, without much fuel needed. The downside is that those trajectories can ...


3

While the overall scope is too broad, let me address the "low energy transfer" part. Firstly, the idea of "gradually increasing apogee" doesn't save you any fuel. Chandrayaan-2 did that because of limited thrust. The only saving to be had here is picking an engine with a slightly lower mass. Secondly, a "ballistic capture", as ...


3

A dynamic system, with at least 3 massive bodies, will have chaos that can, in theory, be exploited to reach (almost) arbitrary positions within said system at close to zero $\Delta v$ over very long time spans. This is the "Interplanetary Transport Network". This sounds very alluring, but it's easy to be misled into believing this has much ...


3

Transfer Regardless of how escape from the Earth system is achieved, orbital mechanics pose some restrictions on travel time. The minimum velocity transfer possible is an elliptical transfer orbit touching Earth's orbit when closest to the Sun, and Jupiter's orbit when farthest from the Sun. The transfer time is then 940 to 1060 days, depending on where in ...


2

There's two ways to view your proposal; If you can do impulse burns but not constant acceleration burns, like your mention of multi-stage accelerations suggests, then what you describe is a series of Hohmann transfers into higher orbital altitude, but instead of doing an orbital circularization burn, you continue with another apogee rising burn as the ...


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