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This post contains a link to Dropbox.
The file turns out to be a text file of the Visual Basic version of GW Johnson's XYTRAJ.BAS program.
https://www.dropbox.com/scl/fi/w6i6xswi … wt8uz&dl=0
and! Here is the zipped up EXE file:
https://www.dropbox.com/scl/fi/9jrp31um … j0p9a&dl=0
Ignore the Dropbox ads and just use the download icon in the upper right.
Move the file from Download to the folder where you want it, and double click the Zip file icon.
After the program appears you can double click on it and it will open it's little DOS window.
GW Johnson is thinking about writing documentation for the program.
The above link was tested on a second Windows 10 machine and this time it worked.
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For GW Johnson re discussion in Rocket Engine Design topic.
Thank you for once again summarizing the issues that must be addressed when designing a rocket engine.
I added an index entry in Post #2 of the topic, in an effort to try to improve the usability of the forum for knowledge storage.
The forum has been primarily a venue for interactions between participants with a very short shelf life. The knowledge captured is then lost and nearly impossible to find.
The questions RGClark has been asking are ideal for providing a framework for education in a complex subject.
In the most recent exchange, it seems to me that RGClark may be asking about rocket engine design, and your answer appears to reflect your experience with ram jets.
In a recent post in another topic, you seemed to hint that there might have been experiments done with a needle shaped probe that was inserted into a nozzle throat to vary the area of the throat through during flight. If you will add a post to the Rocket Engine Design topic, I will create an index entry to take readers right to that post.
A related question is how inserting such a needle into the throat of a running rocket engine would affect the sea level / vacuum issues you described in your recent post.
RGClark has inquired about how injecting fluid into the intake might affect the sea level / vacuum conundrum. I would be surprised if the idea would have any effect, but it might cause turbulence and that would be interesting.
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For GW Johnson!
I got the patent for that.
Let's look for that patent next Sunday. I assume it is assigned to the corporation, or the customer?
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That patent was for adapting a simple showerhead injector whose ports were the gas generator throat for a fixed flow design, to the presence of a throttle valve upstream as the gas generator throat in a variable-flow system capable of arbitrary fuel flow commands.
I had to maintain choked flow at the ports of the injector while avoiding flow acceleration on the way to the next ports. The fixed-flow form had a very subsonic Mach number in the injector. For the variable flow form I had to step the ID at each set of ports to maintain a constant high internal Mach near 0.8 to 0.9. The port sizes had to be large enough not to unchoke the throttle valve at its max area, but also small enough to still choke at the min valve area. And it had to do this despite the intensely 3-D supersonic expansions and shock-downs just downstream of the throttle valve throat.
That throttle worked as a side-inserted pintle into a throat blast tube. I played a key role in making that throttle valve item work, too, to include working out the real-world ballistics of a solid propellant device with such a variable throat, which showed why linearized throttle control logic systems always failed (those motors invariably blew up). Those ballistics were unbelievably non-linear in nature. Once we had the nonlinear adaptive gain determined, we never lost another motor!
My injector patent was assigned to Hercules Aerospace, where I worked at the time in their McGregor plant (it had been Rocketdyne Solid Rocket Division when I first went to work there). This thing was verified extensively in live fire ground tests and was ready to fly when the McGregor plant was closed. The JV partner ARC put it into their Coyote drone for the Navy. It was originally intended for a ramjet upgrade to the AIM-120 AMRAAM. USAF never fielded that upgrade (and there is a rather sordid story behind that failure), although the Europeans have fielded their version of the very same idea: Meteor.
-- GW
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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For GW Johnson re #554
Thanks for additional details about that patent!
***
This post is about a discovery I made about the blogspot.com site where you've been building up a library of articles.
You did not reveal that the site has a search capability, similar to the one that NewMars offers.
You've been telling people to deal with the complexity of your index structure that very few folks are going to have the patience to deal with.
https://newmars.com/forums/viewtopic.ph … 03#p232603
In the post above I report on discovery of the search tool.
I hope you will begin to learn how to use that search tool to quickly find articles by including a tag in the title.
The tag could be as simple as a date of the form YYYYMMDD
I'll bet that if you include a tag like that in the title, you can enter a search in the search window for that tag, and the blogspot software will take you there directly, instead of having to fool around with that convoluted index structure.
Quoting from the article:
This one is not that much reduced in payload capability (6.7% vs 7.5% for the Edge-of-SOTA SSTO and the TSTO). It increased its launch mass a little, being 1487 metric tons, vs 1325 for the edge-of-SOTA SSTO and 1401 for the TSTO. Yet they are all 3 in the same basic class of vehicle sizes. I did select 9 engines, so the diameter is valid, and the L/D is “good”. There is no reason the more modest hydrogen engine technology might not serve, and serve well.
In this scenario, the payload delivered to orbit appears to be 200,000 Kg. (100 metric tons)
The vehicle itself is delivered to orbit along with the payload, so if it is used as construction material it is not a lost investment.
My observation is that if the payload is a single passenger with luggage in a life support capsule, then the difference between the mass of that payload and the computed 200,000 kg could be used for reusability.
In that case, the cost of that one person delivery would be fuel, plus amortization of the vehicle plus maintenance costs.
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For GW Johnson...
kbd512 has posted a comparison of LH2/LOX for SSTO with CH4/LOX
https://newmars.com/forums/viewtopic.ph … 21#p232621
kbd512 appears to be offering the substantial increase in energy density of RP1 over LH2 as a compelling argument in favor of the carbon alternative to LH2. I ** think ** the mass of the vehicle may be in play here.
If you can find a few moments before Sunday's meeting, please revisit your calculations that showed that the carbon based solution would fail for the SSTO application. In your report on the study, you indicated that you stopped working on the carbon solution when you saw that it could not be made to work.
My guess is that kbd512 is missing something important, but from my perspective his argument about saving mass on the vehicle by using heavier propellant is interesting. Again from my perspective, there must be a tradeoff at work, and it would be both interesting and helpful to be able to see how those two curves interact.
It is possible you have already answered the question with graphs that explicitly show the curves of interest.
For all...
GW's work posted in the exrocketman.blogspot.com site is now available by a single click.
https://exrocketman.blogspot.com/search?q=ssto
The paper is detailed, with plenty of graphs to illustrate the text.
I am guessing here, but it is possible that anyone offering an opinion without reading the paper is indeed likely to be missing something important.
*** For GW Johnson... it seems clear that in composing his reply, kbd512 is not taking engine performance into account. The evidence he presents seems ( as I read it ) to be based upon the greater density of carbon based fuel. It appears (again as I attempt to understand the argument) that the mass of the vehicle will be reduced by using a fuel with greater density. Your argument appears to focus on engine performance, which I gather is substantially poorer than for LH2/LOX.
I am hoping it might be possible to show all these elements in a single graph, so that the superiority of LH2/LOX over any of the carbon based solutions becomes clear.
Here is a short list of factors in play as I understand the situation:
1) mass of fuel
2) mass of vehicle
3) performance of engines
4) ISP of fuel
5) drag at the face of the vehicle (assuming no drag after face)
6) gravity pull is equally applied with the lighter vehicle the winner
? other factor I'm missing
Update: https://exrocketman.blogspot.com/search?q=11032024
The revised one-click URL is working, and since it is a unique string, it will deliver just the article you are looking for!
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For GW Johnson re inert mass vs fuel mass...
In the ongoing discussion of LH2 vs carbon fuels, we have history available.
kbd512 has reminded us recently that the Delta IV family of rockets used an LH2 first stage. The family was successful in the sense that it met US military requirements over several decades, but it could not compete in the commercial arena. For the purposes of discussion of the advantages and disadvantages of LH2 vs carbon fuels, the Delta IV history provides useful information, as kbd512 has shown in a post comparing one member of that family to a SpaceX vehicle.
In your analysis of LH2 vs carbon fuels for an SSTO, the results seemed to favor LH2, all factors taken into account.
This is unlike other competing analysis which does NOT take all factors into account.
A detail that is important in any analysis is the proportion of inert mass to fuel. I asked Google to look up the inert mass figures for the Delta IV family of rockets, and it showed inert mass that I deduce must represent the very best that Boeing and United Launch Aliance could achieve while simultaneously meeting their customer objectives.
Inert mass refers to the rocket's structural weight, including the tanks, engines, and other components, minus the fuel.
For the Delta IV family of rockets, the approximate proportions are:
Common Booster Core (CBC) (First Stage):
Propellant Mass: 200,400 kg (441,800 lb)
Empty Mass (Inert Mass): 26,760 kg (58,990 lb)
The inert mass to fuel ratio is roughly 1:7.5, or approximately 13.3% inert mass to fuel.
Delta Cryogenic Second Stage (DCSS):
Propellant Mass: 27,220 kg (60,010 lb) (5-meter version)
Empty Mass (Inert Mass): 3,490 kg (7,690 lb) (estimated from Delta 4H data)
The inert mass to fuel ratio is roughly 1:7.8, or approximately 12.6% inert mass to fuel.
GEM 60 Solid Rocket Boosters:
Propellant Mass: 29,789 kg (65,673 lb) (Gross mass minus empty mass)
Empty Mass (Inert Mass): 3,849 kg (8,485 lb)
The inert mass to fuel ratio is roughly 1:7.7, or approximately 13% inert mass to fuel.
Important Notes:
Variations exist within the Delta IV family depending on the specific configuration (e.g., Delta IV Medium, Medium+, Heavy).Propellant mass fraction (ratio of propellant mass to total stage mass) is commonly used to discuss rocket efficiency. A higher propellant mass fraction means less inert mass per unit of fuel, indicating a more efficient design.
The provided data is based on available information, and slight variations may occur due to different sources or calculations.
More details on specific configurations can be found in sources like Wikipedia.
Delta IV Heavy - Wikipedia
The Delta IV Heavy (Delta 9250H) was an expendable heavy-lift launch vehicle, the largest type of the Delta IV family. It had the highest capacity of any operat...
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WikipediaDelta 4M+(4,2)
Delta 4M+(4,2) ... American orbital launch vehicle. As Delta 4 medium but with 2 x GEM-60 solid rocket boosters and a 4 m diameter payload fairing. AKA: Delta I...
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Encyclopedia AstronauticaDelta 4M
Delta 4M. ... American orbital launch vehicle. Basic Delta-4 vehicle with no strap-ons, the core vehicle, and RL10B-1 upper stage with a 4 m diameter payload fa...
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Encyclopedia Astronautica
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For GW Johnson....
I asked Google the same question about the Falcon 9 family of rockets.
What is interesting (to me for sure) is that the Falcon 9 first stage is reusable, but that does not seem to be much of a factor in the inert mass fraction.
Mass Ratios | Glenn Research Center | NASA
The Falcon 9 rocket's inert mass (structure, engines, avionics, etc.) is a small fraction of its total liftoff mass, with the majority being propellant (fuel and oxidizer). While specific numbers vary slightly between Falcon 9 versions and mission profiles, a general proportion is around 5-7% for the first stage and 3% for the second stage for inert mass to total mass at liftoff. This means that roughly 93-95% of the rocket's mass at liftoff is propellant.
Here's a more detailed breakdown:
First Stage:
The first stage of the Falcon 9 Block 5, for example, has an empty mass of around 22,200 kg, while the total mass at liftoff (including propellant) is 433,100 kg. This results in a mass fraction of about 0.05, or 5%.Second Stage:
The second stage has a much smaller dry mass (around 4,000 kg) and a smaller propellant load compared to the first stage. This leads to a higher mass fraction, but the overall propellant-to-inert mass ratio is still heavily skewed towards propellant.Overall:
The Falcon 9's overall structure mass fraction is relatively low, around 5% for the first stage, and even lower for the second stage, highlighting the rocket's high propellant mass. This is a key factor in its ability to deliver substantial payloads to orbit.
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Just to try to round out the story a bit, here is information about the time tested Proton rocket:
Per Google:
Search Labs | AI Overview
Based on the information available, the inert mass fraction for the Russian Proton rocket, specifically the Proton-M variant, can be calculated using the provided stage data
.
The inert mass fraction is defined as the inert mass of a stage divided by its total mass (inert mass + propellant mass). Here's how it breaks down for each stage of the Proton-M:First stage:
Empty mass: 30,600 kg (67,500 lb)
Propellant mass: 428,300 kg (944,200 lb)
Total mass: 30,600 kg + 428,300 kg = 458,900 kg
Inert mass fraction: 30,600 kg / 458,900 kg = ~0.067 or 6.7%Second stage:
Empty mass: 11,000 kg (24,000 lb)
Propellant mass: 157,300 kg (346,800 lb)
Total mass: 11,000 kg + 157,300 kg = 168,300 kg
Inert mass fraction: 11,000 kg / 168,300 kg = ~0.065 or 6.5%Third stage:
Empty mass: 3,500 kg (7,700 lb)
Propellant mass: 46,562 kg (102,652 lb)
Total mass: 3,500 kg + 46,562 kg = 50,062 kg
Inert mass fraction: 3,500 kg / 50,062 kg = ~0.070 or 7.0%Fourth stage (optional) – Briz-M:
Empty mass: 2,370 kg (5,220 lb)
Propellant mass: 19,800 kg (43,700 lb)
Total mass: 2,370 kg + 19,800 kg = 22,170 kg
Inert mass fraction: 2,370 kg / 22,170 kg = ~0.107 or 10.7%The inert mass fraction varies between different stages of the Proton rocket, ranging from approximately 6.5% to 10.7%. The Block DM (fourth stage) is another potential fourth stage with an inert mass fraction of approximately 14% (2,440 kg inert mass / 17,490 kg total mass).
It is important to note that the inert mass fraction for rockets generally falls within a range of 3% to 15%, depending on the vehicle and propellant combination. Smaller rockets may have a slightly higher inert mass fraction due to the relationship between propellant volume and tank surface area.
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For GW Johnson re Electron rocket...
https://www.google.com/search?q=rocket+ … client=img
This is the web site we were looking at during the Google Meeting
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For kbd512 re Google Meeting discussion:
Here is a link to the Comparing Propellant Combinations document:
https://www.dropbox.com/scl/fi/44rtyihd … rnz5x&dl=0
GW might be willing to prepare another document like this one that is tailored to the specific requirements of the carbon SSTO project.
For kbd512 ... here is the image of fuel comparison:
https://www.dropbox.com/scl/fi/kuynpxiw … hh0ww&dl=0
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For GW Johnson...
Thank you for offering to guide us through the iterative process to develop a Real Universe qualified plan for a Carbon SSTO.
Please think about how we might approach this.
I understand it will be an iterative process, and that computers can help but they can't yet do the whole job.
We should be able to document each stage, and give our members and readers the opportunity to follow along.
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