The Space Plane Corporation

PhotonBytes · 2025-01-13 06:21:22

space-plane.org

Updates to our progress:
https://photonbytes.com/2025/01/13/how- … rporation/
https://photonbytes.com/2024/12/20/the- … simulated/

Our 250 metric ton lifting hull and shock rider design uses a combine cycle air breathing rocket. Our recent simulation break though yielded 17 metric tons of useful payload to orbit. It the Isp numbers were based on two white papers, one attached in the link here:

https://drive.google.com/file/d/16jeWe2 … sp=sharing

Another idea of mine unrelated to the above is building a lighter than air structure, that is a sky scraper filled with lighter than air gas (not hydrogen). That can be used as a space-gun to launch objects to escape velocity:

https://photonbytes.com/2023/05/10/ligh … km-height/
DALL%C2%B7E-2024-01-17-07.17.13-A-vertical-30km-tall-cylindrical-building-with-a-2.5km-radius-constructed-from-lightweight-materials-and-filled-with-lighter-than-air-gas.-The-build-171x300.png

tahanson43206 · 2025-01-13 08:37:54

This post is reserved for an index to posts that may be contributed by NewMars members over time.

The animation produced by this company inspires the hope that the vision may become reality in the Real Universe.

I am looking for Real Universe achievements to be reported in this topic.

Ma Nature will pass judgement on the validity of the concepts in play.

Best wishes for success in turning vision into reality.

(th)

PhotonBytes · 2025-01-22 00:46:28

Recent Update:
Latest Simulation shows more like 5-10 tons to LEO because we need the deorbit burn to return to Earth.

https://photonbytes.com/2025/01/22/500k … rporation/

Last edited by PhotonBytes (2025-01-22 00:47:06)

PhotonBytes · 2025-07-19 12:30:31

We now have 3 ways to take off with our spaceplane:
https://www.facebook.com/share/v/1AepgmNZDw/
1. Standard
2. Refueling truck chase down the runway (personally I think this one is a joke, might as well do the catapult thing like on carriers but on a runway but what do I know? Obviously I didn't come up with this option but I ran the numbers for it anyways.)
3. AN-225 Air Drop from 36,000 ft at Mach 0.8

Each with increasing useful payload capacity as you go down the list from 7, 12 to 19 tons of useful payload.

Last edited by PhotonBytes (2025-07-19 12:31:11)

RGClark · 2025-07-24 11:37:49

I do think a combined-cycle airbreathing+rocket vehicle can work as an SSTO, however there is a key gap in your argument. Looking at the article you cite:

Binbin Lin, Hongliang Pan*, Lei Shi and Jinying Ye
Effect of Primary Rocket Jet on Thermodynamic
Cycle of RBCC in Ejector Mode
Int J Turbo Jet Eng 2017; aop
https://drive.google.com/file/d/16jeWe2 … hP__9/view

it requires scramjet propulsion. But no one has gotten a scramjet to work for more than a few seconds of positive net thrust.

Perhaps you should aim first for ramjet only needed for the upper limits of the airbreathing part of the flight to ca. Mach 5 to 6.

This is the approach aimed for by Skylon. Skylon has been reborn by the way:

INVICTUS: Europe’s Advanced Hypersonic Test Platform by ESA/Frazer-Nash.
https://www.youtube.com/watch?v=zwMHvsNCmQE

I am inviting contributions to my LinkedIn group on SSTO here:

SSTO - Single Stage to Orbit.
https://www.linkedin.com/groups/13205030/

Bob Clark

Calliban · 2025-07-24 15:27:42

The work done by GW Johnson has shown that at altitudes higher than 100,000' and speed greater than Mach 6, air breathing engines become ineffective. The density of the air is too low to produce much thrust at any achievable compression ratio. And shock heating becomes more of a problem the faster the vehicle travels.

This suggests to me that a space plane would work better as the lower part of a TSTO. Imagine a plane that can carry an upper stage within a payload bay, which is released at 100,000'. That altitute is only one tenth what is required for a stable orbit and velocity is only about 30% of orbital velocity. But the relative velocity means that the upper stage engines have high propulsive efficiency and the vacuum at 100,000' allows efficient expansion. So the upper stage should reach orbit with a good mass ratio, which is more conducive to reusability.

tahanson43206 · 2025-07-25 10:04:22

For PhotonBytes ...

Your vision of a space plane remains of interest to members of this forum. It was initially presented as a Blender drawing.

https://newmars.com/forums/viewtopic.ph … 46#p205246

You have reported progress in development of plans for the idea to move from vision to reality.

It's not clear (to me for sure) if you and your collaborators are making progress.

The vehicle you showed us would appear to have potential as the third stage of a traditional rocket.

Your writings seem to suggest you and your collaborators are thinking of trying to fly SSTO.

Are you interested in having your ideas reviewed in the context of the education program GW Johnson is currently developing?

Do you already have someone reviewing your ideas?

(th)

PhotonBytes · 2025-07-26 07:40:10

The core vision is SSTO. That is a total of 9 tons of deliverables to orbit: 1 ton of people, 3 tons of deorbit fuel and 5 tons of useful payload to orbit. TSTO ideas are thought experiments to boost our payload beyond 5 tons and are secondary to SSTO which is our core objective. We wouldn't even entertain the idea of 3rd stage. Most we would concede is a suborbital vehicle if our numbers fall short. But we're not there yet so SSTO is still our primary objective.

Have a look in our files section of our website for reference material about the latest research by NASA and China for burns beyond mach 6 and 100,000ft. We are currently doing 2D CFD engine analysis and could use some help to confirm our design capabilities

Files:
http://space-plane.org/files.htm
Videos :
https://www.facebook.com/share/1AetL2Y59C/

Help wanted:

1. Combustion Specialist

We are currently looking for a combustion specialist with expertise in computational fluid dynamics (CFD) to run internal flow simulations and assist us with our rocket based combined cycle (RBCC) engine design.

2. Control Expert

We need a control specialist to fix phugoid (pitch) oscillations. You can see this in some of the older Facebook videos when the space plane is not simply fixed to 11 deg pitch.This is a very common problem and we need a control expert to fix this. There might be significant propellant saving in doing so, this will again boost useful payload beyond 5 tons without the need to resort to TSTO.

Last edited by PhotonBytes (2025-07-26 09:36:28)

PhotonBytes · 2025-07-27 07:14:41

This is the file to read on the very first tab there are
White paper references for beyond mach 6 and 100,000ft
http://space-plane.org/files.htm
2024.02.01 - Specific impulse and thrust for Mach numbers from 0.3 to 25. Array export to binary file via VBA macro: XLS - 89 kB.

PhotonBytes · 2025-07-27 08:32:42

https://www.facebook.com/share/v/15emv9oTFG/
Video commentary on 2D CFD simulation currently underway.

Last edited by PhotonBytes (2025-07-27 08:42:15)

SpaceNut · 2025-07-27 14:29:12

For scramjet propulsion (Mach 6 to 10)
Air-Breathing Engines: Scramjet engines are a type of jet engine that utilizes atmospheric oxygen as an oxidizer for combustion, unlike rockets that carry their own oxygen supply.
Efficiency: This feature makes scramjets potentially more efficient for hypersonic flight within the atmosphere, as they don't need to carry the extra weight of an onboard oxidizer tank.
Altitude Limit: The reliance on atmospheric oxygen limits scramjet operations to altitudes where sufficient oxygen exists to sustain combustion.
Challenges at High Altitudes: While scramjets are highly efficient at high Mach numbers, the density of atmospheric oxygen decreases with altitude, which will eventually limit their operational ceiling. The maximum operational altitude for scramjets will depend on various factors like engine design, fuel type, and desired performance characteristics

So from the altitude at mach 6 to orbit we must have a source that is stored onboard to achieve the mach 10. That is the amount of time to achieve for a given thrust and burn rate of onboard fuel that remains from the climb to altitude which used the earths atmosphere.

PhotonBytes · 2025-07-28 05:25:24

Yes there is liquid hydrogen as well as liquid oxygen stored. After scramjet we switch to air breathing rocket mode then finally rocket mode. On take off we do the reverse:

1. Rocket (Uses onboard LOX)
2. Air breathing rocket(ejector rockets)
3. Scramjet

Rocket based Combined Cycle
https://trace.tennessee.edu/utk_gradthes/842/

Check out this video and check out the in video telemetry (also in spreadsheet link) and the column for total mass.

https://www.facebook.com/share/p/16jZdBBdDx/

That number will drop from 250 to roughly 85 tons by burning mostly liquid oxygen and some hydrogen. Most of the weight about 165 tons is the former since it's heavier atomically. Whatever remains in orbit minus 1-3 tons of deorbit fuel mass and 1 ton of crew and passengers and the agreed upon weight of the chassis(updated now to 70 instead of 75 tons) is useful payload. Basically anything above 72-74 tons. The later being more conservative as it will be harder and take longer to deorbit with just 1 ton of fuel remaining.

SpaceNut wrote:

For scramjet propulsion (Mach 6 to 10)
Air-Breathing Engines: Scramjet engines are a type of jet engine that utilizes atmospheric oxygen as an oxidizer for combustion, unlike rockets that carry their own oxygen supply.
Efficiency: This feature makes scramjets potentially more efficient for hypersonic flight within the atmosphere, as they don't need to carry the extra weight of an onboard oxidizer tank.
Altitude Limit: The reliance on atmospheric oxygen limits scramjet operations to altitudes where sufficient oxygen exists to sustain combustion.
Challenges at High Altitudes: While scramjets are highly efficient at high Mach numbers, the density of atmospheric oxygen decreases with altitude, which will eventually limit their operational ceiling. The maximum operational altitude for scramjets will depend on various factors like engine design, fuel type, and desired performance characteristics
So from the altitude at mach 6 to orbit we must have a source that is stored onboard to achieve the mach 10. That is the amount of time to achieve for a given thrust and burn rate of onboard fuel that remains from the climb to altitude which used the earths atmosphere.

Last edited by PhotonBytes (2025-07-28 09:19:40)

GW Johnson · 2025-07-28 09:57:20

It sounds like you are talking about a combined-cycle ramjet-rocket design of some sort. Don't confuse ramjet with scramjet. Only the external compression and capture features of the inlet are the same. Everything about scramjet downstream of capture is geometrically incompatible with ramjet. Your outfit might want to talk to an old retired guy like me. I did a lot of rocket and ramjet work, some air turborocket work, even a tad of pulse detonation work. I also designed high speed vehicles, and did hypersonic heat transfer. I might be of some help as a consultant.

GW

Last edited by GW Johnson (2025-07-28 09:57:49)

PhotonBytes · 2025-07-28 19:41:53

Update on scramjet tech:
We do not plan to run a traditional scramjet, but instead add fuel (H2) rich exhaust from the ejectors to incoming supersonic air that will further combust and augment thrust. This fuel rich ejector exhaust should also make the overall combustion process more stable, since it is pre-burned.

GW Johnson · 2025-07-29 09:14:16

Photonbytes:

I do not yet understand what this "airbreathing rocket" is, or your overall propulsion concept. But I gather from your postings that its performance is based on CFD predictions.

I am aware that CFD can do marvelous things these days, but also that it still can only predict from the models built into it. Combustion is not yet fully modeled by anybody, except by "assuming the answer", that being a burn based on mixture in the cell. So the CFD predictions can still be way wrong! The gold standard for CFD propulsion models is still actual test with physical hardware.

I'm no expert in CFD codes, but I do know about that circular logic fallacy when it comes to combustion models. Too many people blindly trust the computer these days, when they really need to be open-minded skeptics. It works pretty good for external aerodynamics, even at entry conditions. Not so well regarding combustion, especially at extreme conditions.

As I said in the other post, I'm an old retired guy, but still able to do a bit of consulting. I started out in the slide rule days. I did rockets, ramjet, air turborocket, pulse detonation, and some other things, plus vehicle aerodynamics and flight dynamics, heat transfer (even hypersonic), some stress-strain, and a whole lot of other things, too.

GW

Last edited by GW Johnson (2025-07-29 09:18:55)

Calliban · 2025-07-30 02:12:26

The latest results from Venus Aerospace.
https://www.nextbigfuture.com/2025/07/h … lanes.html

Impressive technology. It is described as a ramjet engine, that uses a rocket to achieve an intial subsonic boost. As airspeed increases into the Mach range, it transitions into a detonation engine. At high Mach speeds, the velocity of incoming air will exceed the velocity of shockwaves travelling through air. So detonations can provide useful thrust, as they can only travel backwards through the combustion chamber. It isn't clear to me whether this is pulsed power or not. Detonation requires premixed fuel air mixtures. But if the incoming air is already hot, then combustion will naturally occur as increased speed. But it cannot occur more rapidly than the rate of mixing. Hydrogen is the natural fuel for such a detonation engine, because its high molecular speed allows for very rapid diffusion of fuel molecules into air. It also supports a very high flame speed for the same reason. Lighter molecules travel faster.

Vibration from these engines may be a stumbling block. Detonation results in higher peak chamber pressures. If it is pulsed, then vibration will severely shake the airframe. Also, at Mach 10, the temperature and force acting on the compression inlet nozzles will be extreme. This clearly isn't something that could make it all of the way to orbit in a single stage. But it would be a very effective lower stage. If it can exit the sensible atmosphere on a suborbital trajectory, then a second stage could be released from a paload bay. If material and structural problems are sumountable.

Dissociation is not an easy problem to solve. It happens when energetic molecules at the high end of the boltzmann tail break apart, usually into oppositely charged components. There is a work function associated with dissociation, because it pulls apart counter electrically charged species. You don't get the energy back without cooling the combustion products, which allows them to recombine. That only occurs after the exhaust exits the engine. Dissociation robs energy from the system. But it does also reduce exhaust molecular weight. But not enough to be useful.

Last edited by Calliban (2025-07-30 02:29:36)

PhotonBytes · 2025-07-30 09:22:56

Some papers you might want to consider:
https://www.sciencedirect.com/science/a … 611160068X
http://dx.doi.org/10.1016/j.energy.2015.08.017
https://arc.aiaa.org/doi/10.2514/6.2015-3610

And this Rocketdyne video:
https://www.facebook.com/share/v/19wJ76GqXk/

Everything we do is experimental, and CFD is a start. A lot of advancements have occurred in the last 2 decades in computing power (GPU) and CFD.

GW Johnson wrote:

Photonbytes:
I do not yet understand what this "airbreathing rocket" is, or your overall propulsion concept. But I gather from your postings that its performance is based on CFD predictions.
I am aware that CFD can do marvelous things these days, but also that it still can only predict from the models built into it. Combustion is not yet fully modeled by anybody, except by "assuming the answer", that being a burn based on mixture in the cell. So the CFD predictions can still be way wrong! The gold standard for CFD propulsion models is still actual test with physical hardware.
I'm no expert in CFD codes, but I do know about that circular logic fallacy when it comes to combustion models. Too many people blindly trust the computer these days, when they really need to be open-minded skeptics. It works pretty good for external aerodynamics, even at entry conditions. Not so well regarding combustion, especially at extreme conditions.
As I said in the other post, I'm an old retired guy, but still able to do a bit of consulting. I started out in the slide rule days. I did rockets, ramjet, air turborocket, pulse detonation, and some other things, plus vehicle aerodynamics and flight dynamics, heat transfer (even hypersonic), some stress-strain, and a whole lot of other things, too.
GW

Last edited by PhotonBytes (2025-07-30 09:36:56)

tahanson43206 · 2025-08-03 19:48:48

For PhotonBytes...

GW is working on a spreadsheet to calculate a complete SSTO design.

The prototype has run through scenarios with LH2, methane and RP1.

He is looking for ** real ** data on two specifics and perhaps your company can help:

a) mounts for engines which do NOT gimbal
b) mounts for engines with gimbal
3) plumbing to carry fluids from tanks to engines

The spreadsheet depends upon real data to produce reasonable results.

(th)

PhotonBytes · 2025-08-05 09:10:00

Are gimbals for engines for the purpose of thrust vectoring?

Is this a multi fuel design? Or still deciding between purely one or the other fuel type for all engines?

Can you list all the engines and mention the fuel type, isp and thrust for each?

Why not use only h2/o2 for superior isp for all engines? It's tough enough already to get into orbit with h2 what's the idea of also considering methane and RP1? Is it size? You want smaller but heavier fuel tanks vs larger but lighter ones?

RP1:

H2 vs RP1 isp chart for all engine types:
https://share.google/images/d27WFuAidwtJezxAc

We looked into RP1!

RP1 is only good for suborbital craft such as passenger or military strike craft. Add a kerosene rocket at the back of LOX external fuel tanks for say an F-15 and you got your self a suborbital fighter craft that can jump from New York to Tokyo in 30 minutes. In this case it makes sense since all military jets already use jet fuel so all they need is a rocket engine compatible with it and liquid oxygen. Jets have wing pylons that can carry large fuel tanks so just put the Lox in that with a small rocket engine built into the back if it! Then you have twin rockets for the F-15 to fly half way round the world turning fighter pilots into sub-orbital astronauts. With twin rockets the jet can even turn and maneuver.

But not for SSTO

But it's a downgrade to consider replacing any hydrogen with RP1 for an SSTO craft for any engine That's the conclusion we reached years ago. Keep it all h2lLOX was what we learned to max out overall isp since getting into orbit is all about isp.

All engine types relevant can be modified to use it so there's no obvious reason to switch out of h2. Both the rocket and space plane equation show that a lower isp will result in a lower propellant mass fraction not structurally possible.

With the recent advance in rotating detonation jet engines it's even more incentive to use hydrogen for mach 10 possibilities with the jets alone.

Or am I missing something?

Supporting references:

Spaceplane equation
http://space-plane.org/docs/calc/Space- … uation.pdf

Others:
https://drive.google.com/file/d/1EH6DOV … p=drivesdk
https://drive.google.com/file/d/1EL7ZeB … p=drivesdk
https://apps.dtic.mil/sti/citations/ADA467749

tahanson43206 wrote:

For PhotonBytes...
GW is working on a spreadsheet to calculate a complete SSTO design.
The prototype has run through scenarios with LH2, methane and RP1.
He is looking for ** real ** data on two specifics and perhaps your company can help:
a) mounts for engines which do NOT gimbal
b) mounts for engines with gimbal
3) plumbing to carry fluids from tanks to engines
The spreadsheet depends upon real data to produce reasonable results.
(th)

Last edited by PhotonBytes (2025-08-05 12:53:27)

SpaceNut · 2025-08-06 14:38:35

tahanson43206 wrote:

For PhotonBytes...
GW is working on a spreadsheet to calculate a complete SSTO design.
The prototype has run through scenarios with LH2, methane and RP1.
He is looking for ** real ** data on two specifics and perhaps your company can help:
a) mounts for engines which do NOT gimbal
b) mounts for engines with gimbal
3) plumbing to carry fluids from tanks to engines
The spreadsheet depends upon real data to produce reasonable results.
(th)

These question have been looked up for the Delta rocket and Atlas family of rockets

PhotonBytes · 2025-08-07 09:59:22

Is he designing a vertical launch SSTO rocket or a horizontal take off SSTO spaceplane?

SpaceNut wrote:

tahanson43206 wrote:
For PhotonBytes...
GW is working on a spreadsheet to calculate a complete SSTO design.
The prototype has run through scenarios with LH2, methane and RP1.
He is looking for ** real ** data on two specifics and perhaps your company can help:
a) mounts for engines which do NOT gimbal
b) mounts for engines with gimbal
3) plumbing to carry fluids from tanks to engines
The spreadsheet depends upon real data to produce reasonable results.
(th)
These question have been looked up for the Delta rocket and Atlas family of rockets

Last edited by PhotonBytes (2025-08-07 11:58:24)

tahanson43206 · 2025-08-07 12:00:10

PhotonBytes... please look for correspondence in the PhotonBytes Postings topic. The purpose of the topic is to avoid placing an ID over yours in the active list.

In this post: https://newmars.com/forums/viewtopic.ph … 67#p233367
you asked for clarification of the nature of the SSTO spreadsheet that GW Johnson is developing.

The current version is for rocket vertical SSTO expendable. No such rockets exist.

No spaceplanes exist either.

My theory on this is that since space plane is so much harder than straight up, let's make sure we can get straight up working.

Your space plane ** must ** have engine mounts, so the mass of those would be a useful input to the straight up spreadsheet.

If you want to see a spreadsheet for space plane, you can help by assisting with development of the walk-before-run spreadsheet.

A straight up expendable does not have mass allocated to heat shields, wheels or other niceties that are needed for space plane.

GW's preliminary results indicate that if inert mass is at or under 4% of lift off mass, then the straight up model will reach LEO with a small payload other than the rocket itself.

I am interested in the straight up SSTO because is provides a way to deliver ready-to-live-in habitats to an orbital hotel.

The engines can be returned to Earth in a special purpose return vehicle, which would (of course) be designed for reusability.

If you can design a space plane with 4% inert mass, then indications are it would work.

(th)

PhotonBytes · 2025-08-07 21:09:59

https://www.nasa.gov/centers-and-facili … a-success/
https://images.nasa.gov/details/PN068%2 … 09-27-2023
https://ntrs.nasa.gov/citations/20250000643

Rotating detonation engines will make SSTO even easier.

RGClark · 2025-08-11 07:54:54

PhotonBytes wrote:

…
Why not use only h2/o2 for superior isp for all engines? It's tough enough already to get into orbit with h2 what's the idea of also considering methane and RP1? Is it size? You want smaller but heavier fuel tanks vs larger but lighter ones?
RP1:
H2 vs RP1 isp chart for all engine types:
https://share.google/images/d27WFuAidwtJezxAc
We looked into RP1!
RP1 is only good for suborbital craft such as passenger or military strike craft. Add a kerosene rocket at the back of LOX external fuel tanks for say an F-15 and you got your self a suborbital fighter craft that can jump from New York to Tokyo in 30 minutes. In this case it makes sense since all military jets already use jet fuel so all they need is a rocket engine compatible with it and liquid oxygen. Jets have wing pylons that can carry large fuel tanks so just put the Lox in that with a small rocket engine built into the back if it! Then you have twin rockets for the F-15 to fly half way round the world turning fighter pilots into sub-orbital astronauts. With twin rockets the jet can even turn and maneuver.
But not for SSTO
But it's a downgrade to consider replacing any hydrogen with RP1 for an SSTO craft for any engine That's the conclusion we reached years ago. Keep it all h2lLOX was what we learned to max out overall isp since getting into orbit is all about isp.
…

It had been thought that hydrolox was the best propellant to use for a rocket SSTO because of its high ISP. But closer analysis revealed that dense propellants are better for the purpose because their higher density results in lower tank mass, a significant component of the vehicle dry mass.

See discussion here:

Alternate Propellants for SSTO Launchers.
Dr. Bruce Dunn
Adapted from a Presentation at:
Space Access 96
Phoenix Arizona
April 25 - 27, 1996
https://forum.nasaspaceflight.com/index … ach=587468

The tankage for hydrolox is about 3 times heavier, based on density, than kerolox for a rocket SSTO. But for an airbreather it’s even worse: about 10 times heavier!

Bob Clark

GW Johnson · 2025-08-11 08:15:59

I show a similar result, in terms of the "tank efficiency" R-factor. R = Wp/(Wp + inert), basically that fraction of the loaded tank that is actually propellant mass. The tank must contain the volume of the propellant, and larger is inherently a heavier tank, just because it has a larger set of surface areas. But if the density is low, that tank does not contain very much actual mass. For tank section L/D's in the neighborhood of 5:1, with stainless steel balloon construction, I'm showing overall (fuel and oxidizer tanks combined) R's near only 0.94 with LOX-LH2, when its nearer 0.98 or 0.97 with both LOX-RP1 and LOX-LCH4. The oxidizer-to-fuel ratio "r" figures into that, as well as densities of oxidizer and fuel. I was throwing in small allowances for a few frames and stringers in my tank inert masses, plus leaving the upper dome unfilled of liquid, in order to have a vapor pressurization space.

The effect of higher Isp with hydrogen is quite dramatic, probably the largest influence, but the tank efficiency effect is also significant, as it directly increases your stage inert mass fraction. Very quickly, you run out of payload fraction as you try to impose a large dV upon your stage with lower-Isp propellants. Hydrogen does not buy you as much as its Isp might suggest, but that higher Isp effect does offer a bit more payload fraction, despite the higher tank inert.

The problem with all this is the VERY strongly-constrained nature of the stage design problem: the vehicle as launched must be a low drag "clean" shape of the right L/D, or you must start doubling your estimate of the drag loss to cover.

The thrust/weight at liftoff has to be high enough for the vehicle to reach significant subsonic speed while still at very low altitude, in order not to burn up the majority of its propellant before it ever hits the speed of sound. Otherwise, you will be doubling or even tripling your expected gravity loss you have to cover.

Empirically, that liftoff acceleration must be about half a gee net, above weight, or T/W = 1.5 at liftoff. But, the actual engines as sized to create that thrust level still have to fit behind the stage, or else you end up having to double your drag loss that you must cover.

Increasing those losses drives up the dV you must demand of the stage quite dramatically, which drives up mass ratio and propellant mass fraction. At any given inert fraction, you run out of payload fraction VERY quickly! And lower Isp makes that worse even quicker!

You see the mathematical problem here: this is a VERY strongly-constrained optimization, not a free optimization for which you can just run one simple rocket equation calculation.

GW

Last edited by GW Johnson (2025-08-11 08:37:15)

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