Have you noticed, 3I/ATLAS is well and truly on its way out of the Solar System? It has afterall passed through its closest approach to the Sun (perihelion) and is well on its way to a close encounter with Jupiter in mid-March – given this, it would seem to be the ideal moment to contemplate a mission to catch it up.
Many readers will immediately object at this point since hasn’t it conclusively been shown by such papers as this one, that a spacecraft mission from Earth is completely infeasible?
That may be the case but the reader will have overlooked a significant and important detail if they believe so. Most of these previous papers have assumed direct transfer to the target in question – 3I/ATLAS – and have not addressed (to any depth at least) the possibility of indirect missions.
This is where I come in with my extremely powerful software development ‘Optimum Interplanetary Trajectory Software‘ (OITS), a tool I designed to study precisely the indirect possibility – i.e. gravitational assists (or sling-shots) and/or ‘Oberth Manoeuvres’ along the way to the target – and it can handle the direct case also. What amazing software, hey?
You may well have heard of the gravity assist (GA) but what on Earth is an ‘Oberth Manoeuvre’?
This is where a spacecraft under the gravitational influence of a massive body (in this case the Sun) waits to achieve the closest approach (periapsis, or perihelion for the Sun) and then applies thrust at this most propitious point to achieve a high heliocentric speed, in this case resulting in it being shot out of the Solar System, and towards the interstellar object which will have travelled a huge distance by this time.
To summarise the results, there is a way to achieve intercept using a ‘Solar Oberth’ but launch would have to be in the year 2035 to allow optimal alignment between Earth/Jupiter and 3I/ATLAS, and the flight duration would be 50 years, but this could be reduced marginally.
I provide an animation created by OITS of just such a trajectory.
My quest to number-crunch the Mars (Aldrin) Cycler is progressing with every iteration of my software adding more and more encounters, be they Earth or Mars, each adding weight to the calculations.
For instance look at the plot below.
There are a full 80 encounters (horizontal axis) with the cumulative thrust DeltaV plotted on the vertical axis.
A least-squares fit to this yields a straight line with gradient equal to 528.18 m/s per encounter.
This gives an estimate of what would be needed of this Mars Cycler over a long period of time, if the Cycler is to be mantained.
I’ve been progressing with my research into a Mars Cycler where the orbits of Earth and Mars are NOT assumed to be circular NOR coplanar, in other words the orbital paths followed are realistic rather than theoretical.
The following plot shows the cumulative DeltaV thrust achieved with each encounter, where there are 53 encounters altogether (more to come).
The straight line best fit is also provided and the gradient of this straight line is 531.94 m/s per encounter.
This is actually a very useful parameter as it informs us of what would be needed of the proplusion system by the cycling spacecraft over a long period of time.
I’ve been working on the ‘Mars Cycler’, a problem which is normally addressed by assuming the Earth and Mars are in circular, coplanar orbits.
However, I’ve been wondering: ‘What happens in the real case, where Mars has a slight inclination to the Earth’s ecliptic plane, and in fact both Earth and Mars follow slightly elliptical paths?’
Well the answer is that a spacecraft must then use propellant to ensure the Mars Cycler is maintained. But how much propellant?
I can quantify this in terms of DeltaV, using some software (based on OITS) I have developed in C++ expressly for this purpose, the problem being that you need a long string of Earth/Mars encounters to get a good handle of what is happening.
Look at the attached plot which starts in May 2029 and has 13 consecutive encounters. It provides a cumulative DeltaV as the spacecraft visits each planet in turn.
By the way, you might want to see a Mars Cycler at work on my YouTube Channel here:
I’ve been doing a little algebra. Let me state the problem.
Let us say we have a swarm of space sails flying edge on to the interstellar medium (ISM). This swarm lies in a plane at right angles to its velocity relative to this ISM.
Now lets bring in an element of the unknown, a small object, X, of non-negligible mass makes its presence felt, not in the form of an image taken by some of the space sails (thus we have no precise idea as to the location of X as the swarm flies by it), but its presence is felt by its gravitational influence on several nearest members of the swarm of space sails.
The question I have been asking is this: ‘Can we work out from its gravitational influence on the space sails, the consequent mass of the unknown body X?’
The answer to this question requires some thought as to in WHAT WAYS could the trajectory of the space sails be altered by X, and therefore X make its presence felt.
Having contemplated this for some time, I discovered 3 measurable parameters of a space sail’s trajectory which could help in this regard, these are listed below. Think about this for a while and then see if you too can come up with any of these (or possibly ones I hadn’t considered).
(1) The change in speed of the space sails. Thus as the sails approach the body, they will accelerate as they are initially attracted by it and then decelerate back to their original speed, after they have flown past the point of closest approach.
(2) There will be a change in direction of motion of each of the swarm members affected, in other words they shall be deflected by a certain angle, which is measurable. I suspect that this would be more obvious than any perturbation in speed mentioned by (1) above and would provide the better estimate of the mass, particularly when the swarm’s speed is high (20% of the speed of light for example).
(3) Finally the least obvious one. As all the space sails fly past the mysterious X, they will be deflected as mentioned in (2) by a certain angle in a CERTAIN DIRECTION, and this direction will be different for each one. Another way of stating this is that the velocity before the encounter and after encounter will occupy a common plane, and the position of the body X WILL ALSO LIE IN THIS PLANE. Thus by calculating this plane for each space sail in turn, and then determining the point at which all these planes intersect, we are able to determine the precise position of X and from this we can, with the help of (1) and (2), better calculate the mass of X.
I should say I’ve derived an algorithm which exploits (2) & (3) of the parameters mentioned above and uses least-squares fits to generate the mass of an object as a swarm of sails flyby it. It’s working quite well.
I have been considering the feasibility of a mission to 1I/’Oumuamua (under the banner of Project Lyra of course) in the context of a SpaceX Starship launch vehicle.
The word ‘Starship’ here is a contraction – as it is often used to express the combination of Super-Heavy first stage, with Starship second stage.
To be precise, I assume Starship Block 3, which (according to wiki at least) will have its maiden flight next year, and is distinguished from the Blocks 1 & 2 by the use of upgraded Raptor 3 engines (amongst other things) as opposed to the existing Raptor 2, to significantly enhance the overall performance of the launch vehicle.
The nub of the issue for Project Lyra, is what should the demands on the launch vehicle be? In other words, what orbit should the Starship target precisely?
The answer to this question is a minefield for any mission designer; it is not simply a matter of choosing the interplanetary trajectory with minimum overall ΔV (in other words minimum velocity increment from the onboard propulsion system), but actually there is a complex interplay between the launcher specification, its capabilities, the exact nature of any additional stages, and the overall aim of the mission – in this case to catch ‘Oumuamua.
However there is one fundamental, inescapable requirement of any Project Lyra mission powered by chemical rockets, and that is the spacecraft MUST encounter Jupiter, whether that be a powered gravitational assist (GA) – where the thrust is applied at perijove (otherwise known as a ‘Jupiter Oberth Manoeuvre’, JOM) – or a passive Jupiter encounter which is just a straight forward GA, without any rocket thrust.
With this in mind, the first port-of-call for any investigation requires answering the question ‘can Starship inject a payload into an escape mission to Jupiter?’
Unfortunately the Super-Heavy/Starship combination was not optimized with injection directly into an Earth escape orbit (or ‘hyperbolic’ orbit) in mind. Its methalox combination (i.e. methane and liquid oxygen), is designed perfectly to carry hugely massive payloads into LEO (Low Earth Orbit) with ease and reliability. It is from this LEO that refuelling of the Starship can take place, from which the rejuvenated upper stage can inject into any escape orbit with ease, to virtually any destination in the Solar System actually, and this is where Elon Musk’s Mars colonization plans enter the stage.
The cost of this plan is financial. There would have to be on the order of 10 Starship tankers loaded with propellant launched into LEO, before only one single Starship could carry the required propellant mass.
Now with all this in mind, is the Starship indeed capable of a launch to Jupiter? The characteristic energy C3, a measure of the energy required of the launch vehicle, would have to be at least around 100 km2s-2 for a mission to Jupiter. That turns out to be very difficult for the Starship, even with additional rocket stages attached to the spacecraft inside the cargo bay.
There is however an old trick for a mission designer which comes in very handy when the problem is reaching a distant target (like Jupiter in this case) whilst using a much smaller C3 than that for a direct flight to Jupiter. This is known as a ‘V∞ Leveraging Manoeuvre’ (VLM).
A VLM reframes the target needed by the launch vehicle and moves it from the planet in question, in other words Jupiter, around ~ 5.2 au from the Sun, to a point in deep space not quite so far away, hence the reduction in C3. A Deep Space Manoeuvre (DSM) is then conducted, followed by a return to Earth where an encounter takes place, and having significantly increased its energy as a result of this encounter, then the spacecraft flies towards Jupiter without much difficulty.
There are 4 different types of VLM, and these are shown in Table 1.
The first is a return to Earth after 1 year, i.e. with n=1, requiring the least C3, then comes n=2 with a much higher C3 , then n=3, and finally the n=4 option which requires the highest C3 of all. Trajectories n=2,3, & 4 are summarized for you in Figure 1, where they are all depicted in their entirety, though in practice it is likely that only one would be used per mission. The n=1 option would obliterate the view of Earth’s orbit so is not shown in this Figure 1.
Having carried out some preliminary analysis into which VLM from Table 1 would be optimal for Project Lyra and Starship, I have found that probably the best option is the first, i.e. the 1 year resonance, since any C3 higher than 0 km2s-2, is extremely challenging for a direct mission by a Starship Block 3.
My remit when I performed this research was to ignore the option of in-orbit refuelling, even though it has a huge potential. Thus it would seem it is to the least challenging of those options in Table 1 we must refer, if we are to achieve any kind of payload to ‘Oumuamua. The next question is ‘which combination of extra rocket stages do we attach to the Project Lyra payload in order to achieve a C3 = 0 km2s-2?’
Figure 1
There are several possible options from which we can choose and let me first point out that I haven’t tried ALL permutations of stages, that would entail a far more detailed analysis than I have performed here and would be worthy of a paper of itself. However, let us just take it as read for the moment that we shall use two stages, the first a cryogenic combination of liquid propellants LH2/LOX, i.e. the Centaur V rocket stage (which is in production) and also as a second stage, the STAR 48B solid propellant booster. The specification and the ΔV budget for this combination with a 500 kg payload is shown in Table 2.
It so happens I have studied the combination of Centaur and STAR 48B to get to ‘Oumuamua in this paper here, however this was exploiting a NASA SLS rocket and NOT a Starship, and also adopted an older version of the Centaur.
So if one assumes a Centaur V and a STAR 48B with a 500 kg Lyra probe on top (the same mass as the New Horizons spacecraft), then the Starship Block 3 cannot ON ITS OWN, inject this combined payload into an escape orbit with C3 = 1 km2s-2. (Note here that since the required C3 = 0 km2s-2 happens to be a PARABOLIC ESCAPE ORBIT w.r.t Earth, which introduces mathematical singularities, I decided to marginally increase the required performance to C3 = 1 km2s-2, with a perigee of 175 km altitude. This is because these orbital elements allow the target to be slightly HYPERBOLIC rather than PARABOLIC, which doesn’t alter the demands on launcher performance that much but does simplify the calculations.)
I invite you now to look at the results of the optimization process in Figures 2-4. This optimization includes the possibility of incorporating any coast arcs, though none were actually deployed.
Figure 2
Figure 3
Figure 4
Although the Starship cannot inject the payload and combination of stages itemized in Table 2 into the required escape orbit, can we still use some of the thrust from the Centaur V as well? Let us delve a little deeper into this possibility.
First of all, note in practice that for an Earth escape mission, the launcher would most likely inject into an intermediate parking orbit around the Earth, before the Centaur completed its burn to Earth escape. However bear in mind that this is a feasibility study, so a direct ascent to escape was assumed, to get a handle of what the Starship is capable of.
A Centaur V has a restartable rocket engine, meaning the thrust can be effectively turned off at any time during its burn, and more than that can be turned on again after a period of coast. This allows the Centaur V to achieve the required C3 of the target orbit after the Starship has done its best, and there will still be more of the Centaur’s LH2/LOX propellant left-over and unused. Would this spare propellant go wasted then?
If we examine Figure 2, we find that there is a remaining ΔV from the Centaur V of ~6.75 km/s. Normally a mission designer would seek to use ALL this available propellant at Earth escape since delaying would result in undesirable and very considerable boil-off and leakage of the cryogenic propellants which the Centaur exploits. But there is room for hope, go here for an investigation of how a Centaur might be restarted after a considerable elapsed time, using COLD (Cryogenic Operation for Long Duration) technologies. The authors are quite optimistic that such technology could be developed and implemented with numerous benefits for space missions.
It is at this point we turn to my ‘Optimum Interplanetary Trajectory Software’ (OITS), which can solve the over-arching problem of which interplanetary trajectory is best for catching up with ‘Oumuamua, given that the initial C3 = 1 km2s-2 at Earth.
So I proceeded with my investigations using OITS and found the trajectory in Table 3.
This is depicted in Figure 5. The plan in question requires two burns after Earth escape: the first at a Return to Earth (~6.6 km/s) and the second at the Jupiter encounter (3.6 km/s).
The encounter at Earth (an ‘Earth Oberth Manoeuvre’, EOM) exploits a ΔV equivalent to that remaining in the Centaur V after the Earth escape with ~150 m/s subtracted. This deduction is to account for 4% reduction in propellant over the course of 1 year after injection, due to boil-off and leakage, which is a low figure but apparently within the capabilities of the COLD technologies elaborated in the aforementioned paper.
This leaves the Jupiter Oberth Manoeuvre, JOM, requiring a ΔV of exactly 3.6 km/s which turns out to be precisely the amount available from the STAR 48B, allowing the 500 kg payload to be accelerated to ‘Oumuamua after 43 years from launch.
Figure 5
So is this the best plan for Project Lyra and SpaceX? It seems not necessarily.
If we reduce the Project Lyra payload to 100 kg, then there IS another way!
The alternative plan is to forget the enormous, super-powerful Starship and use a SpaceX Falcon Heavy Expendable instead. Go to a previous blog of mine, here, which shows that there is an architecture which can reach ‘Oumuamua after only 28 years.
Yet a caveat is in order here. Such is the complexity of the problem, there being SO MANY permutations of trajectories and upper stages (and launchers for that matter), I am convinced there are better alternatives available for Project Lyra, and it is my long-term aim to discover exactly what they are.
[1] Hyperbolic Excess at Earth, V∞, where C3 = V∞2
On the request of a colleague, I have solved the problem of exploiting the powerful SpaceX Starship (in fact the yet to be launched Block 3 variant) to lift a spacecraft so that it can catch up with 1I/’Oumuamua, the now rapidly receding ‘interstellar object’.
This object sped through the inner group of Solar System planets in 2017, and then headed away again, back into interstellar space.
It left a mountain of questions in its wake, which a spacecraft flyby mission could answer, returning tremendous scientific data in the process.
I used the currently in production Centaur V cryogenic liquid propellant stage and also the STAR 48B solid propellant booster, in addition to the Starship upper stage.
To do this research I needed to optimize both the geocentric launcher-ascent-to-orbit trajectory (i.e. the trajectory which maximises the mass-to-Earth-escape) and the subsequent heliocentric interplanetary trajectory designed to chase ‘Oumuamua down, and into which the Starship inserts the Lyra spacecraft.
The geocentric and heliocentric trajectories are shown in the plots below. The first three show the Starship ascent directly to an Earth escape orbit, and the fourth shows the interplanetary trajectory to catch ‘Oumuamua after 40 years.
My software ‘Optimum Interplanetary Trajectory Software’ (OITS) was used in the process of conducting this research.
I have had many queries concerning the calculation of the likelihood of 3I/ATLAS’s orbital plane lying within 5° of that of the ecliptic of the Solar System.
This calculation appears in the paper ‘Is the Interstellar Object 3I/ATLAS Alien Technology?’ which can be found here.
The calculation exploits a simple equation based on the ratio of two areas on a spherical surface.
If you imagine the plane of 3I/ATLAS’s orbit has a vector perpendicular to it, and the inclination of its orbit therefore is the angle between this vector (more precisely known as the ‘angular momentum’ vector) and the axis representing the ecliptic north pole.
The probability therefore that the inclination of the orbit lies within 5° of this ecliptic north pole, is the area, A, of a circular cap surrounding this North pole (extending 5 degrees either side of this pole) divided by the total area of the sphere, S.
Now by Archimedes ‘Hat Box’ theorem we have A=2πr2( 1 – cos(5°)), and we have the total area of a sphere, S = 4πr2. The ratio A/S =( 1 – cos(5°))/2 = 2e-3 or ~ 0.2 %. This latter figure is quoted in the paper. In fact to be exact we need to multiply this by 2, since we might be within 5° of the south polar axis also.
There is an alternative logic, however, which has been put to me which comes up with a much higher probability. This I believe to have a flawed logic. I shall elaborate below.
The counter-argument goes like this. Let’s say we represent the ISOs as darts and assume they arrive, isotropically (meaning from all directions with equal likelihood) at the Solar System which is represented again as a perfect sphere.
The idea is that the number of darts, L, arriving within 5° latitude of the ecliptic plane should then be divided by the total number of darts, N, and thus the probability of the plane having an inclination < 5° is simply the ratio of these two numbers, i.e. L/N.
What one find when one does this calculation is that you get a MUCH HIGHER PROBABILITY, than that derived in the paper. In other words L/N >> A/S.
So why the disagreement?
The reason is because there is a fallacy in the argument of the latter scenario. Let’s say that a dart arrives at 4° latitude above the ecliptic, does it necessarily follow that the inclination of its orbit is also exactly 4°?
It turns out: not at all! If the velocity when it hits the sphere is vertical (so parallel to the ecliptic north) then, its inclination will be exactly 90°, even though it strikes the sphere at only 4° latitude!
Thus by adding up all the darts arriving at less than 5° latitude we shall arrive at a spurious answer since we shall be unwittingly including in our calculations, darts that have a much higher orbital inclination than 5°.
Human existence, and for that matter all life, has been a story of death. Observe that life attempts to preserve itself by defying death.
In fact evolution is simply an attempt by the universe to learn more about itself.
In my view the universe has set death a purpose. It is teaching life to know and understand.
Look at humanity.
People have died for generations upon generations. But scientific understanding and technology have always advanced.
Particularly since around the time of the industrial revolution, we have begun to outwit death, our longevity has been increasing, child mortality rates have plummeted. This has largely been a consequence of deeper scientific understanding, particularly in medicine.
Death catches up with us all, has all this loss of life been futile? Has every living thing struggled in vain?
Not at all. By individuals dying, life is collectively knowing more and more about the universe which enables us to circumvent death.
What wonders will our successors discover? Maybe a way of defying death forever?
Perhaps even the ultimate celebration of life, a kind of unification of all life that has ever existed?
A colour contour plot showing direct missions to object 2014 UN271, otherwise known as ‘BB’ in honour of its discoverers Bernardinelli and Bernstein. This comet is huge (~ 135 km) and very distant, yet it turns out we can launch a mission.
The coloured bits represent viable direct missions to BB -in particular, the deep blue bits would be achievable by a sufficiently powerful launch vehicle. The blank areas represent infeasible combinations and occupy most of the plot. Note there are yearly opportunities up until a launch in around 2031.
If we are going to realise a direct mission to flyby BB, then it’ll have to have a combination of launch Date/Flight Duration in one of the coloured areas shown.
Alternatively, we could send a probe on an indirect trajectory as indicated by my animation, below. The software used to generate this was Optimum Interplanetary Trajectory Software (OITS) devised and created by me.
Inidrect method of reaching comet 2014 UN271 solved by OITS
The paper, written by me and Marshall Eubanks, was published by the Journal of Spacecraft and Rockets, here. For the preprint go to arXiv, here, or otherwise refer to my ResearchGate profile here.
By the way my YouTube channel with lots of mission trajectory animations (including Project Lyra) can be found here.
Below is a plot that expresses how much ‘DeltaV’ or in other words, ‘change in velocity’ would be needed to be applied to the asteroid Apophis, at different times along its path, so as to send it careering catastrophically into the Earth on Friday the 13th of April 2029, instead of missing it by around 31,000 km, as it is very likely to do. The Apocalypse Plot is attached.
If you are concerned for my sanity at this point, please do not be; there is method in the madness, go to my article for Principium, ‘Rise of the Serpent God’, here:
Alternatively read the article below:
The best trajectory a SpaceX Starship could take to budge Apophis off course is animated below:
I’ve been investigating the Double Jupiter Gravitational Assist (DJGA) and comparing it with the ordinary Single Jupiter Gravitational Assist (SJGA) which we know well and is propounded in the Interstellar Probe Concept Report. For context, refer to my poster:
I seems that at DJGA (which I invented) is not all that it was cracked up to be.
For your information, please find in the attached plots below:
(a) a passive DJGA (at both encounters) – compared with a passive SJGA.
(b) a passive DJGA at the first encounter followed by a DeltaV=3km/s at the second encounter – compared with an SJGA with DeltaV=3km/s at the only Jupiter encounter.
(c) a passive DJGA at the first encounter followed by a DeltaV=9km/s at the second encounter – compared with an SJGA with DeltaV=9km/s at the only Jupiter encounter.
For further explanation (b) & (c) represent the hyperbolic excess speed w.r.t. the sun achieved by the respective trajectories against the latitude of the escape asymptote out of the solar system. We find in (b) & (c) above that the SJGA out-performs the DJGA.
(a) represents the achieved hyperbolic excess speed w.r.t. the sun against the characteristic energy C3 at Earth Launch (optimal escape asymptote latitude is assumed). We find in (a), there is only a marginal superiority of the DJGA compared with the SJGA, especially at C3>115km2/s2.
In summary I would say there is no appreciable gain in using a more complicated DJGA, particularly when the SJGA out-performs it in most circumstances.
I am still trying to reconcile the above findings with the efficacy of this trajectory when I studied missions to ‘Oumuamua and Borisov. I may be missing something here!
Just been asked is it likely or even possible that 3I/ATLAS, the latest interstellar object originating from some distant Milky Way location, will perform some kind of manoeuvre on its way to Jupiter to rendezvous with it.
My standard response to such questions is that 3I/ATLAS is a comet.
Yet nevertheless, it can be fun to contemplate such a scenario, and I shall not be dissuaded by anyone from doing so, for whatever reason.
I SHALL exercise my imagination if I so desire, sure it is a flight of fancy, but human psyche finds excellent nourishment from creativity and fantasy, and I loathe any rigid dogma that clamps down on that.
Thus the answer to the question is that the level of non-gravitational accelerations observed in 3I/ATLAS are generally FAR lower than those I derived in my paper with Avi Loeb and Adam Crowl, and so it is unlikely that 3I/ATLAS is attempting to use any low-thrust propulsion method to conduct a rendezvous of the planet.
However, high thrust propulsion is an entirely different matter. Currently 3I/ATLAS will side-step Jupiter by a full 0.36 Sun-Earth distances. This turns out to be exactly the Laplace sphere radius for the Sun/Jupiter system.
Another way of looking at it is that this is the approximate border line between Jupiter having a meaningful influence on 3I/ATLAS’s orbital path and the Sun’s influence taking over.
So if 3I/ATLAS wants to rendezvous with the planet, then it will need to deliver a pretty high thrust VERY SOON to ensure it comes within a reasonable distance of Jupiter to allow it to stay in an orbit bound to the planet.
A rendezvous thrust at Jupiter will also be needed, though as discussed in the paper, 3I/ATLAS can be slowed by AERO-CAPTURE exploiting Jupiter’s very thick atmosphere to decelerate the object to below Jupiter’s escape velocity and stay orbiting Jupiter, thus obviating the object needing any precious propellant to achieve capture.
Yesterday I was trying to gauge the measure of the SpaceX Starship in terms of its ability to launch the Project Lyra spacecraft on its way to its destination.
BTW Project Lyra is the initiative to send a mission to catch-up with very rapidly receding interstellar object 1I/’Oumuamua. So exactly how do we get a probe to intercept this object?
Well I discovered ages ago that whatever mission plan we might adopt; a visit to the planet Jupiter, either as a means-to-an-end, or alternatively to exploit it directly by delivering a Jupiter Oberth Manoeuvre would be vital in achieving the mission goal: 1I/’Oumuamua.
So how does the Starship fare as far as lofting a payload that will eventually reach Jupiter is concerned?
The unfortunate answer is not very well. This answer is important because this super-heavy launch vehicle is designed as a part of Elon Musk’s initiative to colonise the planet Mars, which involves an Earth escape (hyperbolic) orbit just like Project Lyra’s – though obviously the former escape is to Mars and the latter is to Jupiter.
So what’s holding it back?
The straight-forward answer is the choice of Starship’s propellants which are methane and oxygen as opposed to the likes of alternatives like NASA’s Space Launch System (SLS) which employs the much more powerful, and efficient, LOX and LH2 combination of cryogenic propellants.
It seems that Starship’s methane/LOX combination is ideal for inserting a payload into LEO (Low Earth Orbit), from which the Starship can then be refuelled and sent off to any destination within the Solar System without much difficulty, but absolutely pants at injecting a payload directly into an escape orbit from launch.
To be fair, Elon Musk has made it quite clear that Starship was designed with this refuelling strategy in mind, but there is a cost, which one would expect for any decision along these lines.
The cost is indeed financial – it would take on the order of 10 Starship tankers filled with the requisite methane+LOX in the cargo bay to fully refuel ONE Starship located in LEO – the logistics and economics are a tough reality check for anyone who has bought into the glib Mars colonisation rhetoric spouted by Elon.
So where does this leave Project Lyra? Well I shall continue my research into using a Starship launch vehicle, and I shall have to add two or even three extra stages to the Project Lyra craft payload itself.
If you are slightly concerned by this profligate use of rocket stages, have no fear, there is plenty of space in the Starship’s huge cargo bays for these rocket stages + the Lyra craft.
As a consequence of exploring the Solar Oberth option to catch up with 3I/ATLAS I have, using my software development Optimum Interplanetary Trajectory Software (OITS), discovered that a mission exists IN THE FUTURE, with a launch in 10 years time, i.e. in 2035.
The video animation can be found on my YouTube channel here, or alternatively look below.
It involves approaching within a solar-distance of 2 Solar Radii, that’s very close to the Sun indeed, to exploit maximum benefit from the so-called Oberth Effect, where all the thrust is applied at the closest approach to the Sun, perihelion.
By so-doing the spacecraft maximises the increment in kinetic energy and heads off towards the target at a huge speed, in this case in excess of 100 km/s (or 0.03% light speed).
The journey takes around 20 years. As a comparison the Voyager probes were launched in 1977 and are still to a limited extent operational after ~48 years, nearly half a century later.
Look below for my demonstration of the Oberth Effect:
My letter to the president, written a few years after 1I/’Oumuamua was discovered and the Project Lyra work had largely drawn to a conclusion, is now looking rather stale since 3I/ATLAS has made its appearance.
It seems that 3I/ATLAS’s anti-tail (which is a plume pointing towards the Sun), has transformed into an ordinary tail (pointing away from the Sun).
My observation is that if these were jets of gas from a braking manoeuvre, then as the object approached perihelion this plume would point approximately towards the Sun, whereas after perihelion this plume would point away from the Sun.
This would therefore be behaviour constituting evidence of alien technology.
A useful diagram is shown below.
Athough this ‘switch’ was indeed observed for 3I/ATLAS, note it happened BEFORE perihelion.
