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This topic is offered for NewMars members who would like to find posts in future.
Occasionally members create posts that are worth rereading at a later time.
The forum has no mechanism in place to find posts easily.
The Search tool is a powerful tool, but it is only as good as the memory of the user.
In addition, search terms we might try often deliver many unrelated posts we have to sift through.
There is a time honored mechanism that SpaceNut just reminded me to consider: bookmarks
in the physical Real Universe, a "bookmark" may consist of nothing more complex that a slip of paper sticking out from between two pages.
In the digital realm, this topic is offered as a way for each member to store a reminder of where a particularly interesting or valuable post may reside.
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This post is reserved for an index to posts that may be contributed by members.
Since this post must be maintained by moderators, it is unlikely to stay up with the flow of traffic, but it may provide a quick lookup capability to the extent it is maintained.
The ** main ** purpose of this new topic is to enable individual members to create their ** own ** bookmarks.
If those bookmarks are helpful to others, that's a bonus.
Index:
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Bookmark:
Author: GW Johnson
Location: exRocketman.blogspot.com
https://exrocketman.blogspot.com/
Rocket Equation exercise with gravity and drag
This study compares SSTO and TSTO and concludes that SSTO is feasible if the vehicle is left in orbit (ie, expendable).
A reasonable scenario for that case is the material of the rocket is intended to be used after delivery to LEO.
The blogspot site has a search window in the upper left corner. Like the NewMars search tool, it is a crude instrument, but it ** did ** find an article in which SSTO is included.
Interestingly, the article cited is a follow up to the one I was looking for:
Monday, March 11, 2024
More-Refined 1- vs 2-Stage to LEOUpdate 6-8-2025: this article is the continuation of a study originally posted 3 March 2024, titled "Launch to Low Earth Orbit: 1 or 2 Stages?". That study produced some broad recommendations, this one produces actual trends of payload fraction and launch weight vs propellant combination, for the all-expendable TSTO and SSTO scenarios.
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This link takes you to the first article that includes ssto:
https://exrocketman.blogspot.com/search?q=ssto
Update: https://exrocketman.blogspot.com/search?q=11032024
This link takes you to the article itself.
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This bookmark takes you to a lively exchange between RobertDyck and kbd512 about robots in combat
https://newmars.com/forums/viewtopic.ph … 50#p232650
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This bookmark is for one of a great number of interesting posts by Photonbytes
https://newmars.com/forums/viewtopic.ph … 46#p205246
This one shows a proposed SSTO vehicle with reusability features. Because we know that this design is unlikely to fly as an SSTO, it is of historical interest as an example of ambitious thinking. However, the vehicle seems feasible for a TSTO configuration, especially if the first stage is itself reusable.
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kbd512 created a post about electric flight
The key concept is that a new fuel cell is available that can take in RP1 (kerosene or diesel fuel) and produce electricity.
The bottom line is performance in both weight and size and the amount of travel produced is greater with the fuel cell.
https://newmars.com/forums/viewtopic.ph … 64#p232664
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kbd512 created a post about rocket engine design and testing...
The post includes links to a video show rocket engine testing, and a link to a web site about the White Sands test facilities.
https://newmars.com/forums/viewtopic.ph … 82#p232682
In the post, kbd512 attempts to explain his statement, supported by GW Johnson, that thrust increases in a rocket engine as it rises in the atmosphere, even though the mass flow through the nozzle remains the same.
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The post at the link below is about drag. The post considers both aircraft and spacecraft.
GW Johnson on drag...
https://newmars.com/forums/viewtopic.ph … 85#p232685
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This post is another of kbd512's many posts on military subjects...
https://newmars.com/forums/viewtopic.ph … 03#p232703
This one makes a case for 155 mm weapons for all services and all applications with modular propulsion packs.
it also reviews tank design (from several Nations) and considers the advance of drones in modern warfare.
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In this post kbd512 compares the Delta 4 and the Falcon 9
https://newmars.com/forums/viewtopic.ph … 13#p232713
Let's consider the energy efficiency of RP1 vs LH2, purely from an energy consumption per kilogram of useful payload perspective.
D4M = Delta IV Medium (all stages fueled with LH2)
F9B5 = Falcon 9 Block V (all stages fueled with RP1)
This stack exchange discussion reviews the history of why LH2 was chosen for Delta IV, and a comparison of LH2 vs carbon fuels.
https://space.stackexchange.com/questio … rst-stages
A key takeaway is that the mass of a rocket is not the same thing as the volume of the rocket.
The volume of an LH2 rocket will be greater than a carbon vehicle, but the mass should be less for equivalent performance.
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tahanson43206,
The fuel volume of a LH2 fueled rocket and the comparatively poor engine thrust-to-weight ratio of LH2 fueled engines is increasing the dry mass of the rocket while decreasing Total Impulse per unit of vehicle dry mass. Delta IV requires more vehicle dry mass per unit of payload delivered to orbit. If you were to fabricate the Delta IV from Carbon Fiber and use the highest-Isp LH2 engine available, the RS-25, Delta IV would still have a higher dry mass per unit of payload delivered to orbit, as compared to Falcon 9.
What do I mean by that?:
IF you make Delta IV's propellant tanks from Carbon Fiber, lopping 40% off the dry mass of its propellant tanks
AND
IF you use a pair of even higher Isp / higher thrust / lower total mass (compared to RS-68A) RS-25 staged combustion engines
AND
IF you only fire that pair of RS-25 engines in a hard vacuum where their thrust is highest
THEN
You only generate 45,052 Newton-seconds of thrust per unit of vehicle dry mass for this "upgraded" Delta IV
IF you fire Merlin gas generator engines at sea level, where their thrust is lowest
AND
IF you keep the Al-2195 Aluminum alloy propellant tanks of Falcon 9, which are much heavier than Carbon Fiber
THEN
Falcon 9 Block V still generates 46,050 Newton-seconds of thrust per unit of dry vehicle mass
45,052N-s total thrust per kg dry vehicle mass <-- LOX/LH2 staged combustion / CFRP tanks / engines fired in a vacuum
46,050N-s total thrust per kg dry vehicle mass <-- LOX/RP1 gas generator / Aluminum tanks / engines fired at sea level
If you apply the same engine tech (staged combustion) and propellant tank (CFRP) tech upgrades to LOX/RP1, then it will result in even more Total Impulse per unit of vehicle dry mass. Even if you somehow doubled the engine thrust-to-weight of LH2 fueled engines, you will still get more Total Impulse per unit of vehicle dry mass. Until vehicle dry mass approaches zero, there is quite literally nothing you can physically do to a LH2 fueled rocket powered vehicle solution to make it more effective than RP1 in terms of vehicle dry mass.
Why does this matter so much for a SSTO?
For a SSTO, you NEED as many Newton-seconds (N-s) of total thrust generated per kilogram of dry vehicle mass as you can possibly get. The reason you need that is because you're accelerating the entire vehicle's dry mass, along with the useful payload, from a dead stop on the launch pad to orbital velocity. More N-s per kg dry mass is "more better", and in comparison to a TSTO, it's also "more required".
How did I figure that out using basic multiplication and division?
1. I know how much total propellant mass each rocket holds (in kilograms).
2. I know the mass flow rate per second (in kilograms per second) and thrust generated (in Newtons), for the Merlin-1D, RS-25D, and RS-68A engines, both at sea level and in a vacuum.
Note:
All of those engines have published thrust ratings for sea level and vacuum, as determined by empirical testing on specialized test stands located at Earth sea level, which are capable of simulating a vacuum environment. As you previously discovered, the mass flow rate doesn't change, unless throttle position changes, and the nozzle geometry doesn't change, either. Only the engine thrust output changes as atmospheric back-pressure rapidly reduces with increasing altitude. This, in turn affects Isp (thrust generated by a specific mass flow rate).
Absent real mass flow rate and thrust data, when designing "paper rockets", you may use 90% of Vacuum Isp associated with a given engine propellant combination and combustion cycle. For example, the real life RD-180 engine, an Oxidizer-rich staged combustion engine that burns RP1, has a sea level Isp of 311s and a vacuum Isp of 338s. For comparison, Merlin-1D is only 282s (sl) and 311s (vac).
For a practical SSTO, we need an engine with the thrust-to-weight ratio (TWR) of Raptor with the Isp of the RD-180. I think this is doable, because Raptor uses larger pumps to feed LCH4 (half the density of RP1), yet still achieves 200:1 TWR. Therefore, a RP1 fueled staged combustion engine with a 200:1 TWR should be feasible using the same simplification process that Raptor went through as 3D printing enabled complex component integration without bolting together separate parts. For this "paper engine", we would use 90% of RD-180's vacuum Isp, meaning 304.2s as our flight-averaged Isp. Merlin's flight-averaged Isp is only 279.9s. That provides an 8% "fuel consumption reduction", or in actually, about 8% greater Total Impulse (49,734N-S/kg of dry mass) if vehicle mass stays the same.
3. I divided total propellant mass by mass flow rate per second for all engines, which tells me how many seconds of total firing time I get.
4. I multiplied engine thrust (in Newtons) by total firing time (in seconds), to arrive at Newton-seconds (N-s) of Total Impulse (It).
Note:
If this was a variable value, as would be the case for the solid rocket motors I used as a child, then you have to integrate the result. For a liquid fueled rocket engine this value doesn't change unless the throttle setting changes or the thrust changes, so no calculus is required to compute the result at full rated engine output for a given altitude / atmospheric back-pressure (sea level or hard vacuum). For a real rocket flight where altitude continuously changes throughout the flight, you briefly spin-up the turbopumps, briefly throttle down the engines near Max-Q and then throttle back up again after your vehicle is supersonic, and then throttle down towards the end of the burn to avoid over-acceleration of the vehicle, you must integrate the thrust measurements. On top of that, you need to know what the propellant residuals must be to avoid exploding your engine turbopumps as the propellant tanks run dry. Electric motor driven turbopumps can suck the tanks dry. Combustion driven turbopumps cannot.
5. I divided Newton-seconds of Total Impulse by vehicle dry mass, to determine how much total thrust a given propellant combination would produce per unit of vehicle dry mass, both at sea level and in a vacuum.
The end result of this 5-step exercise illustrated how dramatically higher Total Impulse is for RP1 vs LH2, for a given vehicle dry mass.
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