For any given vehicle design, what one assumes for mission delta-vees, vehicle weight statements, course corrections, and landing burn requirements greatly affects the payload that can be carried. The effect is exponential: variation in required mass ratio with changes in delta-vee and exhaust velocity.
This analysis looks at trips from low Earth orbit to direct entry at Mars, and for the return, a direct launch from Mars to a direct entry at Earth. The scope is min-energy Hohmann transfer plus 3 faster trajectories (see ref. 1).
The vehicle under analysis is the 2019 version of the Spacex “Starship” design, as described in ref. 2. The most significant items about that vehicle model are the inert mass and the maximum propellant load. For this study, the vehicle is presumed fully loaded with propellant at Earth departure, and at Mars departure. See also Figure 1. Evaporative losses are ignored.
Since a prototype has yet to fly, the design target inert mass of 120 metric tons is presumed as baseline. Uncertainty demands that inert mass growth be investigated. To that end, the average of that design target and the 200 metric ton inert mass of the so-called “Mark 1 prototype” (that average is some 160 metric tons) is used to explore that effect.
As currently proposed, the vehicle has six engines. Three are the sea level version of the “Raptor” engine design, and the other three are vacuum versions of the same engine design (basically just a larger expansion bell). I have already reverse-engineered fairly-realistic performance for these in ref. 3. Because of the smaller bells, the sea level engines gimbal significantly, while the vacuum engines cannot. Thus it is the sea level engines that must be used to land on Mars as well as Earth: gimballing is required for vehicle attitude control.
As shown in Figure 2, the analysis process is not a simple single-operation calculation. The vehicle model provides a weight statement and engine performance. The mission has delta-vee requirements for departure, course correction, and landing, which must be appropriated factored (in order to get mass ratio-effective values). There are two sets of analysis: the outbound leg from Earth to Mars, and the return leg from Mars to Earth.
Each leg analyzes 3 burns. Earth departure, and course correction are done with the vacuum “Raptor” engines, while the landing on Mars is done with the sea level “Raptors” to obtain the necessary gimballing. Mars departure and course correction are done with the vacuum “Raptor” engines (Mars atmospheric pressure is essentially vacuum). The Earth landing is done with the sea level “Raptors” to get the gimballing and to get the atmospheric backpressure capability.
This analysis is best done in a spreadsheet, which then responds instantly to changes in one of the constants (like an inert mass or a delta-vee). That is what I did here.
Referring again to Figure 2, for each burn, there is an appropriate vehicle ignition mass. At departure, it is the ignition mass from the weight statement. For each subsequent burn, it is the previous burn’s burnout mass. Each burn’s burnout mass is its ignition mass divided by the required mass ratio for that burn, in turn figured from that burn’s delta vee and the appropriate exhaust velocity.
For each burn, the change in vehicle mass from ignition to burnout is the propellant mass used for that burn. For the first burn, the propellant remaining (after the burn) is the initial propellant load minus the propellant mass used for that burn. For the subsequent burns, propellant remaining is the previous value of propellant remaining, minus the propellant used for that burn.
After the final burn, the propellant remaining cannot be a negative number! If it is, one reduces the payload number originally input, and does all the calculations again. If this done in a spreadsheet, this update is automatic! Ideally, the propellant remaining should be exactly zero, but for estimating purposes here, a small positive fraction of a ton (out of 1200 tons) is “close enough”.
Thus it is payload that is determined in this analysis. This particular input (payload mass) is revised iteratively until the final burn’s remaining-propellant estimate is essentially zero. That is the maximum payload value feasible for the mission case.
Orbits and the Associated Delta Vees
As indicated in ref. 2, I have looked at a Hohmann min energy transfer orbit, and 3 faster transfers with shorter flight times. All of these are transfer ellipses with their perihelions located at Earth’s orbit. For Hohmann transfer, the apohelion is at Mars’s orbit. For the faster transfers, apohelion is increasingly far beyond Mars’s orbit. Why this is so is explained in the reference. See Figures 3 and 4.
Note that the overall period of the transfer orbit is important for abort purposes. If the period is an exact integer multiple of one Earth year, then Earth will be at the orbit perihelion point simultaneously with anything traveling along that entire transfer orbit. This offers the possibility of aborting the direct entry and descent at Mars, if conditions happen to be bad when the encounter happens. Otherwise, the spacecraft is committed to entry and descent, no matter what.
The cases examined in ref. 1 were all computed for Earth and Mars at their average distances from the sun. The larger transfer ellipse with the longer period occurs when both Earth and Mars are at their farthest distances from the sun. This leads to larger delta vees to reach transfer perihelion velocity for the trip to Mars, and larger velocity on the transfer orbit for the trip back to Earth.
Ref 1 has the required velocities and delta-vees, but the most pertinent data are repeated here:
Transfer E.depdV, km/s trip time, days M. Vint, km/s
Hohmann 3.659 259 5.69
2-yr abort 4.347 128 7.40
No abort 4.859 110 7.36
3-yr abort 5.223 102 6.53
Transfer M.depdV, km/s trip time, days E.Vint, km/s
Hohmann 5.800 259 11.57
2-yr abort 7.548 128 12.26
No abort 7.509 110 12.77
3-yr abort 6.653 102 13.14
I did not examine the worst cases for all the transfer orbits in ref. 1, but I do have the increase in perihelion velocity for the worst case Earth departure on a Hohmann transfer for Mars: 0.20 km/s higher than average. I also have the increase in apohelion velocity for the worst case Mars departure on a Hohmann transfer for Earth: 0.16 km/s higher than average.
I cheated here: I used those worst-case Hohmann increases for all the faster trajectories as well. That’s not “right”, but it should be close enough to see the relative size of the effect of worst case over average conditions. I also used the same additive changes on the entry velocities.
Because of the precision trajectory requirements for direct entry while moving above planetary escape speed, some sort of course correction burn or burns will simply be required. With this kind of analysis, I have no way to evaluate that need. So I just guessed: 0.5 km/s delta-vee capability in terms of propellant reserves.
Because this is just a guess, I did not run any sensitivity analysis on it. However, the delta-vee budget proposed here is factor 2.5 larger than the difference average-to-worst-case for the trip to Mars, which suggests it is “plenty”. It is about factor 3 larger than the difference average-to-worst-case for the return trip to Earth. You can get a qualitative sense of this effect from examining that average-vs-worst case effect.
Propellant Budgets for Direct Landings
With this vehicle (or just about any other vehicle), entry must be made at a shallow angle relative to local horizontal. Down lift is required to avoid bouncing off the atmosphere, since entry interface speed Vint exceeds planetary escape speed. This is true at both Mars and at Earth. Once speed has dropped to about orbit speed, the vehicle must roll to up lift, to keep the trajectory from too-quickly steepening downward.
The hypersonics end at roughly local Mach 3 speeds, which is around 0.7-1 km/s velocity, near 5 km altitude on Mars, and near 45 km altitude on Earth which has about the same air pressure. Up to that point, entry at Mars and Earth look very much alike, excepting the altitude. After that point they diverge sharply, as illustrated in Figure 5.
The descent and landing at Earth require the ship to decelerate to transonic speed, then pull up to a 90-degree angle of attack (AOA, measured relative to the wind vector). Thus, as the trajectory rapidly steepens to vertical, the ship executes a broadside “belly-flop” rather like a skydiver.
At low altitude where the air is much denser, the terminal speed in the “belly-flop” will be well subsonic. I assumed 0.5 Mach, but that might be a little conservative. This is the point where AOA increases to 180 degrees (tail-first), and the landing engines get ignited. From there, touchdown is retropropulsive.
The landing on Mars is quite different. The ship comes out of hypersonics very close to the surface, still at high AOA and still very supersonic. From there, the ship must pitch to higher AOA and pull up, actually ascending back toward 5 km altitude. This ascent is energy management: speed drops rapidly as altitude increases. It’s not quite a “tail slide” maneuver, but it is similar to one.
At the local peak altitude, the ship is moving at about local Mach 1, and pitches to tail first attitude, igniting the landing engines. From there, touchdown is retropropulsive. The Martian “air” at the surface is very thin indeed, as the figure indicates. It may be that thrust is required to assist lift toward bending the trajectory upward: the engines would have to be ignited earlier, and at higher speed, as indicated in the figure. Whether this is necessary is just not yet known.
The low point preceding the local pull-up is at some supersonic speed; I just assumed about local Mach 1.5, as indicated in the figure. That would correspond to a factor 1.5 larger landing delta-vee requirement, implying a larger landing propellant budget.
In either case, I also use an “eyeball” factor of 1.5 upon the kinematic landing delta-vee, to cover gravity loss effects, maneuver requirements, and any hover or near-hover to divert laterally to avoid obstacles.
So, for purposes of this sensitivity analysis, the Earth landing is not of much interest, but the Mars landing is. The sensitivity analysis looks at the effects of Mach 1.5-sized vs Mach 1-sized touchdown delta-vee.
The scope of the sensitivity analysis is illustrated in Figure 6. As indicated earlier, the orbital delta-vee increases worst-case-vs-average, for Hohmann transfer, were applied additively to the departure delta-vees for the faster trajectories. No attempt was made to vary the course correction budgets. Growth in vehicle design inert mass was examined. An increase in the Mars touchdown delta-vee was examined. Nothing else was considered.
The results start with the worst vs average orbital delta-vee sensitivity. These results are given in Figure 7. These are the plots from the spreadsheet, copied and pasted into the figure. There are 4 such plots in the figure: the top two are for the outbound journey Earth to Mars. The bottom two are for the return journey Mars to Earth. Results for all 4 transfer orbit cases are shown simultaneously by using trip time as the abscissa.
Each has 4 data points: these are for the Hohmann transfer at 259 days flight time, the 2-year abort orbit at 128 days, the non-abort orbit at 110 days, and the 3-year abort at 102 days. Be aware that the curves are probably not really straight between the Hohmann orbit and the 2-year abort orbit. I did not run enough fast transfer cases in ref. 1 to get a smooth curve here.
The most significant thing in the left hand figure for the outbound trip is the about-40 ton loss of max payload between average and worst case for the Hohmann transfer. This is a lot less than the about-130 ton payload loss using the 2-year abort orbit instead of Hohmann transfer, or the about-210 ton payload loss for using the 3-year abort orbit.
The average-vs-worst-case deficits are somewhat similar on the faster orbits. The Mars entry interface velocity trend in the right-hand figure is obviously very nonlinear. Yet, all the calculated values fall below the entry velocity from low Earth orbit (LEO). Any heat shield capable of serving for return from LEO will serve this Mars entry purpose, which would be the governing case if the trip were one-way only. There’s only a small change in entry speeds for average-vs-worst orbit case in this estimated analysis.
The return voyage has trends shaped quite differently. For Hohmann transfer, the worst-vs-average payload loss is about 20 tons. The deficits on the faster orbits should be similar. The deficit for using the 2-year abort orbit instead of Hohmann is far larger at about 110 tons, and that’s from a small return payload to begin with.
In the right hand Earth entry interface speed plot, the blue and orange curves in the entry interface plot fall only slightly apart. Note that all the entry velocities are much higher than the just-below-escape speed seen with Apollo returning from the moon. The faster transfer orbits, and even the Hohmann transfer, are substantially more demanding than a lunar return entry. It is clearly the direct-entry Earth return that will size the heat shield design!
Figure 7 – Sensitivity to Worst-Case Orbital Distances vs Averages
The sensitivities to the need for a thrusted pull-up on Mars are given in Figure 9. This follows the same format as Figures 7 and 8. Bear in mind that the nominal design lights the engines for touchdown at about Mach 1 speed. For this analysis, the engines are ignited earlier, at about Mach 1.5 flight speed, to assist lift in pulling up to the Mach 1 “flip”, to tail-first attitude. That makes the landing delta vee about 1.5 times larger. (Note that each case is also factored up by 1.5 further, to cover any maneuver / hover needs for the touchdown.)
What the figure shows is about the same 40-ton payload loss on the voyage to Mars to cover the increased landing propellant requirement for the Hohmann transfer. Effects on the faster transfers are similar. This trend is comparable to the worst-case orbit losses. The return payload is entirely unaffected, as the landing occurs prior to refueling and loading for the trip home.
Both the Earth and Mars entry interface velocities are unaffected by this Mars thrusted pull-up scenario. The orange and blue curves fall right on top of each other.
#1. These results are only approximate! Real 3-body orbital analysis, and real entry-trajectory lifting flight dynamics models, must be used to get better answers. Nevertheless, the trends are quite clear from this approximate analysis.
#2. Flying on faster transfer orbits will cost a lot of payload capability, on both the outbound voyage, and the return voyage. This effect is much worse on the return voyage, where the allowable payload is just inherently smaller.
#3. The effects of worst-case orbital positions-relative-to-average, of Mars and Earth, have a significant effect on payload, but it is only half or less the effect of choosing faster transfer orbits.
#4. The effect of vehicle inert mass growth from the design target of 120 metric tons to an arbitrary but realistic 160 metric tons is comparable to the effect of worst-case vs average orbits on the outbound voyage. However it has catastrophic effects on the return voyage! This is enough to prevent faster-than-Hohmann transfers on the voyage home, for this vehicle model.
#5. The effects of needing a thrusted pull-up for the Mars landing is comparable to the effects of worst-case orbit distances on the outbound voyage. This has no effects upon the return voyage.
#6. It is the direct Earth entry velocity that will design the vehicle heat shield for any vehicle capable of making the return. This is substantially more challenging than was the return from the moon. For deliberately-designed one-way vehicles to Mars, the heat shield design requirements are comparable to entry from low Earth orbit.
#7. My personal opinions are that thrusted pull-up will be needed, along with the need to fly when Earth and Mars orbital distances are worst-case, plus there will be a little inert mass growth (say by 20 metric tons to 140 metric tons vehicle inert mass). That kind of thing is the proper design point for this vehicle, not the most rosy projections! Estimated performance data for this design case (at 140 metric ton inert mass) are in Figure 10 (same basic format as Figures 7, 8, and 9). Note that two of the faster transfers home are precluded. The feasible one has a very small max payload value compared to Hohmann transfer.
Figure 10 – Performance for Worst Orbits, Thrusted Pull-Up, and Some Inert Mass Growth
#8. Bear in mind that the rather high max allowable payload figures feasible to Mars for Hohmann transfer are incompatible with what can be aboard “Starship” for launch to low Earth orbit. The payloads for the faster transfers to Mars look more like what can be ferried up to LEO. That suggests that a faster transfer to Mars is most compatible with the projected “Starship” / “Super Heavy” system design characteristics, as these were evaluated in references 2 and 3.
#9. Bear also in mind that a faster transfer orbit to Mars ought to include abort capability, in case conditions at arrival prove too bad to attempt the landing. There is simply not the propellant available to enter orbit and wait for better conditions. Thus life support supplies must be carried to last the entire period of the transfer orbit, and a full-capability heat shield for direct Earth entry must be used.
#10. The fast transfer home need not be limited by abort capability. It can be a different transfer orbit than the outbound trip. Surprisingly, the shapes of the plotted curves suggest that something faster than the “3-year abort” orbit could be used for the return home.
#11. Given a way to combine two payloads to LEO into one “Starship” by cargo transfer operations on orbit, then (and only then) the very large payloads to Mars indicated for Hohmann transfer become feasible. Like on-orbit cryogenic refueling, this on-orbit cargo transfer capability does not yet exist, not even as a concept (on-orbit refueling at least exists as a concept).
#1. G. W. Johnson, “Interplanetary Trajectories and Requirements”, posted 21 November 2019, this site.
#2. G. W. Johnson, “Reverse-Engineering the 2019 Version of the Spacex “Starship”/ ”Super Heavy” Design”, posted 22 October 2019, this site.
#3. G. W. Johnson, “Reverse-Engineered “Raptor” Engine Performance”, posted 26 September 2019, this site.
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