Upgraded Rocket Hopper as Orbit Taxi

This article is about modifying a pre-existing design rough-out for a suborbital Mars rocket “hopper”,  into a design also capable of operating as a low Mars orbit personnel taxi.  That original rocket hopper design rough-out is covered in the article titled “Rocket Hopper for Mars Planetary Transportation”,  dated 1 November 2023,  and posted on this site.

 

               The Problem

Started with a suborbital “hopper”

               10 persons aboard on p-suits

               Short-term life support plus small luggage

Could it also serve as a low orbit taxi?

               Same payload

 

As indicated in the table just above,  I started with the earlier design rough-out that was only a suborbital “hopper”.  The idea was to carry 10 persons as the payload.  Although the cabin is pressurized,  these persons ride in pressure suits for a safety backup.  There are limited supplies of oxygen and drinking water,  plus minimal snack foods,  for up to a few hours’ ride.  A small luggage allowance was included.  The same payload would be carried to any low orbit destination.

As indicated in Figure 1 just below,  the suborbital trajectory is actually an ellipse in polar coordinates,  one with its periapsis inside the planet.  The vehicle launches into a gravity turn that reaches a suitable velocity and path angle at the entry interface altitude,  coasting from there. 

The best place to do a course correction is the apoapsis outside the sensible atmosphere,  where speeds are lowest and directions are easiest to change.  The entry conditions mirror the exit conditions,  with no burn.  The landing is a direct rocket-braked descent from the end-of-hypersonics point at local Mach 3 (about 0.7 km/s speed). 45 degrees of trajectory “droop” along a straight-line path is presumed.  I factored-up the speed to “kill” by 2,  to budget the final landing mass ratio-effective delta-vee (dV).

As illustrated in Figure 2 just below,  I used a surface-grazing ellipse as the initial transfer trajectory to the 300 km nominal low orbit altitude.  Like the long-range suborbital mission,  the vehicle launches into a gravity turn,  putting it onto the proper path at the entry interface altitude,  at end of launch burn.  Only the path angle is different,  being a lot smaller.  The entry point after de-orbiting is the mirror image. 

There is a small burn at apoapsis to raise the periapsis to the entry interface altitude,  with a period shorter than the target low circular orbit altitude.  This ellipse is the parking orbit in which to “chase” any target in the low circular orbit.  Once synchronized,  there is another small burn to circularize,  followed by a traverse to rendezvous,  plus a budget to actually dock.  Deorbiting is another small burn,  back onto the surface-grazing ellipse that guarantees entry.  The direct rocket-braked landing is identical to that of the long suborbital trajectory,  except that,  as it turned out,  the end-of-hypersonics altitude is higher,  coming back from orbit at the lower entry angle. 

Figure 1 – Suborbital Missions,  Longest-Range Shown  

Figure 2 – The Orbital Mission,  Including “Chase”,  Rendezvous,  and Docking  

To accommodate the more demanding mission,  I resized the candidate LOX-LCH4 engine design,  and revised the inert masses upward a little.  Entry conditions forced me to increase the diameter and length a little,  in order to keep the entry ballistic coefficient down to tolerable values.  The original rough-out had two sets of tanks:  mains and headers.   The landing and course correction propellant was in the headers,  with the launch propellant in the mains.  

This became 3 sets of tanks and two different engine designs.  The launch-and-landing main engines stayed about the same at 30,000 lb thrust,  each of 4,  drawing from the mains for launch and headers for landing.  I was able to increase the expansion ratio and specific impulse a little bit.  See Figure 3 for the basic layout revisions.

But course correction suborbitally,  and all the orbital maneuvering,  rendezvous,  and docking,  really needed much lower thrust levels.  So I sized some lower-pressure,  pressure-fed engines of only 550 lb thrust,  each of 4.  These used a small third set of 800 psi pressurized propellant tanks,  plus a supply of dry nitrogen gas at 2200 psi to power this,  in one of two options examined.  

Figure 3 – Revised Internal Layout at Larger Diameter and Length  

Because the inert fraction increased a bit,  I resized the expansion of the main engines to increase specific impulse a bit,  to compensate as much as possible.  The original “hopper” main engines had an expansion ratio sized for incipient separation at 67% chamber pressure,  if fired in the open air at sea level on Earth.  I raised that to 80%.  See Figure 4 just below.

The idea was to enable easy and relatively inexpensive development testing on Earth.  The change was small,  but every little bit helps.  These being full flow cycle,  turbo-pumped engines of significant thrust level,  I did not want to complicate things by adding vacuum bell extensions that were not regeneratively cooled.  These do not push the state of the art very hard,  being only 2000 psia chamber pressure.

Figure 4 – Reworked Main Engine Design for Slightly Higher Expansion 

The original “hopper” design study convinced me that I did not need the large main engine thrust levels to do course corrections on the suborbital missions,  or orbital maneuvering,  rendezvous,  and docking,  on the orbital mission.  I kept the redundancy of 4 engines,  but sized for crudely only 0.1 gee of vehicle acceleration,  once exoatmospheric. 

Since the propellant quantities would be small,  the simplification of a pressure-fed design would be beneficial.  Alternatively,  since the engines were small,  they could be fed by electric-driven positive-displacement pumps.  Either way,  I picked a simple conical bell shape,  two-piece,  with a bell extension that was not regeneratively cooled,  as shown in Figure 5 just below.  Development testing on Earth could be done without the extension,  but operations on Mars or in space would use the bell extension.  This was not a throttleable design.

I ran numbers both ways for the propellant feed to the maneuvering engines.  I did not like the pumping power required for the positive-displacement pumped option.  It implied very heavy batteries,  even for the modest propellant quantities.  Regulated constant inert gas pressure on the propellant tanks turned out to be the better option.  These used a small third set of 800 psi pressurized propellant tanks,  plus a supply of dry nitrogen gas at 2200 psi to power this.  The chamber pressures were low enough to keep the pressures fairly modest on the tanks,  so that at small size,  they were not that big an inert weight penalty.  See Figure 6 just below. 

There were many false starts and iteration cycles to achieve all of this,  none of which is covered here.  The result is summarized in the unavoidably-busy figure,  Figure 7 just below,  which includes a weight statement that also displays mass fractions.  

Figure 5 – Smaller Maneuvering Engines as Sized   

Figure 6 – Of the Options,  Pressurized Tanks Seemed Best  

Figure 7 – Summary Data for the Final Rough-Out Design 

Of interest would be the various tank volumes.  Bear in mind these are fully filled for the mission to low circular Mars orbit at only low inclinations eastward,  and also fully-filled for the long-range suborbital mission (at low or high inclinations).  The headers and maneuvering tanks are always fully-filled,  but the mains are only partially filled for the shorter-range suborbital missions,  so that entry mass is not too big.

Suborbital ranges from 9400 to just under 500 km were examined in this study.   Their entry angles turned out to be a strong function of the suborbital mission ranges.  All of those suborbital entry angles were considerably steeper than the return from the orbital mission.  They were determined by feasible altitudes at end-of-hypersonics,  and by feasible peak entry gee values.

Figure 8 just below shows the final plots I got of various flight data during entry and final descent.  The Suborbital trajectories form trends,  and the orbital data fall way off those trends.  Upper left is end-of-hypersonics altitude and entry angle vs entry speed.  Upper right are the peak heating values during entry.  Lower left are the trends of peak entry gee,  and average gee during the final rocket-braked landings.  There is a numbered key relating the missions to each data point in each plot.  No gee level exceeds 4.5,  and no stagnation heating level exceeds 12.5 W/cm².  

Figure 8 – Data for Entry,  Descent,  and Landing (E, D, & L)  

Once again in Figure 9 just below,  the suborbital data for surface temperatures also form trends versus entry speed,  with the orbital data falling far off of those trend lines,  plus a numbered mission key.  There are trends for surface temperatures at the stagnation line,  temperatures for its lateral surfaces where flow is still attached,  and temperatures for leeward separated wake zone surfaces.  These were figured for thermal re-radiation exactly balancing convective-only input,  as figured for a “dark” highly-emissive surface,  of thermal emissivity 0.8 as representative. 

Note that with the exception of only the longest-range suborbital mission,  all the rest of these data are under 1600 F,  and would permit exposed-metal construction of 316L stainless steel,  or of Inconel X-750,  or something in that same class!  And that even includes the return from the orbital mission!  Because of the stagnation zone temperature approaching 2000 F on the longest-range suborbital mission,  there needs to be some minimal heat protection in and near the stagnation zone.

In Figure 10 just below,  the format for surface pressures is similar:  trends of suborbital surface pressure vs entry speed at stagnation,  at lateral sides,  and in the separated leeside wake.  The orbital data again fall far off the trend lines.  There is a numbered key to relate missions to individual data points.  Note that no mission,  not even the longest-range suborbital,  exceeds 0.19 atmosphere anywhere.  That would be about 2.79 psi,  very modest indeed. 

Given the hot material strengths reported as part of Figure 9,  that means even a fragile extreme-low-density alumino-silicate ceramic composite could serve as heat protection.  So could ceramic fabric blanket or quilt-type materials,  if they survive wind shear.   Even a thin sheet of 2000 F-capable metal overlying mineral wool insulation would serve,  mounted only locally near the stagnation line. 

Conclusion:  the “hopper” could easily be designed to also serve as a low orbit taxi!

Figure 9 – Trends of Surface Temperatures vs Entry Speed  

Figure 10 – Trends of Surface Pressures vs Entry Speed  

Update 11-22-2023:  The following Figure 11 illustrates exactly why the surface emissivity must be high (a very dark or black surface color,  with a dull finish).  There are exposed metallic construction solutions and a refractory solution with simple alumino-silicate ceramics,  especially away from the stagnation zone,  if emissivity is high.  There ae only ablative solutions available if emissivity is low.

Figure 11 --  More Detailed Hopper/Taxi Heating Data