GW Johnson Postings and @Exrocketman1 YouTube videos

tahanson43206 · 2025-07-01 11:44:35

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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tahanson43206 · 2025-07-02 06:41:37

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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tahanson43206 · 2025-07-02 08:58:14

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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GW Johnson · 2025-07-03 10:39:23

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

tahanson43206 · 2025-07-04 08:07:21

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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tahanson43206 · 2025-07-05 06:21:50

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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tahanson43206 · 2025-07-09 07:18:47

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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Wikipedia
Delta 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 Astronautica
Delta 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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tahanson43206 · 2025-07-09 07:27:01

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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tahanson43206 · 2025-07-09 08:38:40

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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tahanson43206 · 2025-07-13 19:01:54

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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tahanson43206 · 2025-07-13 19:53:23

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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tahanson43206 · 2025-07-13 19:58:48

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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tahanson43206 · 2025-07-18 09:52:27

For GW Johnson re post in SSTO LH2 topic...

Thank you for this strong boost to the topic!

I'll quote a particularly interesting section for a question:

ALL airbreathers (no matter WHAT they are!!!) have a thrust level that is more-or-less proportional to the atmospheric pressure in which they are flying. Up there, 5-to-15 times nothing is still nothing! No significant thrust to push considerable mass! You will neither accelerate nor climb! It is called the "service ceiling" effect. Only rockets are immune to it. Because they do not breathe air.

So why are we limited to just air or just rocket?

It seems to me that there is no reason not to run the two propulsion methods simultaneously.

If you take the Tangent Launch concept (originating with kbd512 but brought back to life by GW Johnson in 2025), it seems to me logical to combine modes as follows:

1) Railroad for 3 kilometers at 2 G to reach just under the speed of sound at sea level
2) Air breathing starts and runs at less than full strength
3) Rocket starts and runs until Air breathing is at full strength
4) Air breathing continues until it is no longer useful
5) Rocket resumes with increasing strength as Air breathing falls off
6) Air breathing stops
7) Rocket continues to orbit

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GW Johnson · 2025-07-19 13:22:07

Adding to the quote in post 563 just above:

To do both airbreathing and rocket, you either do some sort of combined-cycle engine or you do parallel-burn mixed propulsion with both airbreathing and rocket engines aboard.

Combined cycle inherently and significantly compromises the performance of its components, because their geometries are incompatible to one extent or another, but it can occupy most or all of the vehicle frontal blockage cross section area.

Parallel-burn mixed propulsion gets full performance out of its components, but they each occupy only some fraction of the vehicle frontal blockage cross section area, which limits how much thrust each can supply.

For airbreathing, the most mature and ready-to-apply technologies are ramjet and gas turbine. High bypass turbines are subsonic-limited. To go supersonic, you must lower the bypass ratio which increases specific fuel consumption dramatically. To go high supersonic, the bypass ratio is close to nil, and even then most of those are heating-limited to only Mach 2.5, and only reach such speeds in full afterburn, which has horrible specific fuel consumption. There are only a very few gas turbine engines that were ever qualified to fly at Mach 3 to 3.5: those in the XB-70 at Mach 3, those that propelled the SR-71 at Mach 3.2, and the short-life (500 hour), replace-do-not-overhaul turbines in the Mig-25 at Mach 3.5.

The other mature airbreathers are ramjets. These come in two forms: low speed designs, and high-speed designs.

Low speed designs burn from high subsonic to at most about Mach 2, usually only up to about Mach 1.5. They feature pitot/normal-shock nose inlets, and convergent-only nozzles (because in a properly-sized engine, the nozzle throat is unchoked until you reach about Mach 1.05 to 1.1 flight speeds). These usually use V-gutter flame stabilization, and usually require volatile gasoline as fuel. Some sort of colander burner can also work. Combustors and nozzles can be perforated steel double wall, using some of the captured air as cooling air.

High speed designs burn from about Mach 1.8 or higher to about Mach 4 or 6, although no one design will cover that whole speed range. They feature inlets with external supersonic compression features (spikes or ramps), they can be nose, chin, or side entry inlet integrations, and they have convergent-divergent nozzles that are always choked. Up to about Mach 3-to-3.5 in the stratosphere, they can use V-gutter flame stabilization, and they can use perforated-liner air cooling of combustors and nozzles.

Above such speeds, there is no such such thing as "cooling air", and only ablatives and dump-combustor stabilization can be used.

For takeover Mach numbers in the 1.8 to 2.5 range, you need to use modestly-volatile wide-cut fuels like Jet-B or JP-4. For takeover speeds at or above 2.5, you can use kerosene or kerosene-like synthetics, like Jet-A or Jet-A1, JP-5, and RJ-5/Shelldyne-H (a synthetic slightly denser than water).

The nose and chin inlet integrations are low-enough drag to let you fly past Mach 4. The side entry inlet integrations are draggy-enough to restrict you to speeds of Mach 3 to 4.

Anything faster than about Mach 4 presents a severe-to-horrific aeroheating problems, even in the cold stratosphere, something very difficult to solve steady-state, indeed! And the inlet capture features and internal duct heating difficulties are even worse than vehicle nose tip and aerosurface leading edge difficulties, because surfaces able to re-radiate are smaller or completely absent.

GW

Last edited by GW Johnson (2025-07-19 13:30:42)

tahanson43206 · 2025-07-26 18:02:49

For anyone who might have a LinkedIn account, here is a link to an update GW Johnson just applied ...

https://www.linkedin.com/feed/update/ur … 827229696/

Hopefully the link will work for others. That is the URL that was in my address bar as I was looking at the update.

The update is about the cryogenic tank fluid movement concept, and the drawing shows the Patent Pending text we've reported previously.

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SpaceNut · 2025-07-27 14:33:46

yes the counter rotating drum paddle concept to make the fuel flow to the outside under artifical gravity with the center becoming pressurized as the tank isw emptied.

https://media.licdn.com/dms/image/v2/D5 … nyWB6X6RIE

tahanson43206 · 2025-07-29 12:52:00

For GW Johnson!

Thanks for the report of the logout not working as you expect (and have a right to expect) (not always the same thing).

If you have time (and remember it at the time you are logging out) there is a bit of useful information you can request from the forum software.

There two possibilities... one is that the system in not logging you out

The other is that the system is logging you out but not telling you. Those are two ** very ** different things.

You can help the Admins to diagnose the problem if you will scroll down to the bottom of the page to see if your name is listed in the Online section.

If your name is NOT listed, then you are not logged in.

If your name IS listed, then you are still logged in.

There is a timer that will log you out automatically after something like 30 minutes. I'm not sure of the exact time. It may be as little as 15 minutes. Whatever the limit is I hit it frequently when I am distracted while working with the forum. In that case, to recover my work, I open a second login on the same computer, and the forum software allows me to complete whatever I was doing in the other window.

From time to time in the past, we have had members reporting the time out.

When this happens to anyone, just log in with a separate browser session and you'll be able to perform whatever update you had pending.

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tahanson43206 · 2025-08-01 11:21:15

GW Johnson has been working on a number of projects.

Here is a preliminary report on one of them (SSTO related)
https://www.dropbox.com/scl/fi/p6wt1yv9 … vpupf&dl=0

GW has been working on a complex spreadsheet that can (according to initial testing) provide a realistic sizing for components of rockets with carbon and hydrogen fuels, in various payload sizes.

GW asked me to note that while the spreadsheet appears to be working, the inert fraction is still too low to be realistic.

https://exrocketman.blogspot.com/search?q=01082025

The link above will take you directly to the article: Air Launch to Low Earth Orbit

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tahanson43206 · 2025-08-03 12:47:26

For GW Johnson...

The link below is to a file that contains system components needed to run XYTRAJ on your computer.

https://www.dropbox.com/scl/fi/q45r86i3 … cj83e&dl=0

Please ignore the pop up windows from Dropbox, and use the download icon to deliver the file to your download folder.

This file is from Microsoft. It was stored on my computer from about 2010.

It provides the c++ files that the program EXE on your desktop needs.

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SpaceNut · 2025-08-03 18:39:23

tahanson43206 · Yesterday 09:37:42

For GW Johnson re question from PhotonBytes...

https://newmars.com/forums/viewtopic.ph … 94#p233294

Will your current SSTO spreadsheet accommodate hybrid design (Ie, multiple fuels)?

I have tried to clarify that your current work is on non-reusable rockets, and NOT on space planes.

(th)

SpaceNut · Today 14:30:53

AI Overview
The Delta rocket family utilizes gimbaled engines to steer the vehicle during launch. The mass of the gimbal mechanism varies depending on the specific engine and rocket configuration, but it generally contributes a small percentage to the overall vehicle mass, typically around 0.1% to 0.5%. The RS-68A engine on the Delta IV Heavy, for example, has a gimbal system that allows for thrust vectoring, enabling the rocket to steer during ascent. While the exact mass contribution of the gimbal mechanism to the RS-68A engine isn't explicitly stated, it's a relatively small fraction of the engine's overall mass.
Here's a more detailed breakdown:
Delta IV Heavy:
The Delta IV Heavy uses three RS-68A engines, each capable of producing 2.1 million pounds of thrust. The RS-68A is a gimbaled engine, allowing for thrust vectoring for steering.
Engine Mass:
The RS-68A is a large engine, and the gimbal mechanism is a relatively small part of its overall mass. The RL10B, used on the Delta IV upper stage, has a mass of 664 lbs and includes a large carbon-carbon nozzle extension.
Gimbal Contribution:
Studies and examples suggest that gimbals typically contribute between 0.1% and 0.5% to the total launch mass. For example, in the case of the Antares 100 rocket, gimbals on the NK-33 engines contributed about 0.16% of the liftoff mass.
Delta II:
The Delta II used the RS-27 engine, and later the upgraded RS-27A, which also employed a gimbaled system for steering. The RS-27A produced 1,054 kN (237,000 lbf) of thrust.
Other Factors:
The mass of the gimbal mechanism can also be influenced by factors such as the size and complexity of the actuators, hydraulic systems, and other associated components

AI Overview
Delta family of rockets: engine mount mass
Determining the exact mass of engine mounts within the Delta family of rockets can be difficult because the manufacturer often considers this information proprietary, and rarely discloses it as a separate component. However, related data can be analyzed to infer some insights into their mass:
1. Engine dry mass
Delta IV RS-68 Engine: The RS-68 engine used in the Delta IV Common Booster Core has a dry mass of 6,604 kg.
Delta RS-27A Engine: The RS-27A engine used in the 7000 and 8000 series Delta rockets has a dry mass of 1,091 kg.
Delta II RS-27A Engine: The RS-27A engine featured in the Delta II rocket has a dry mass of 1,146 kg.
Delta IV RL10B-2 Engine: The RL10B-2 engine utilized in the Delta IV's second stage has a dry mass of 301 kg.
2. Structural considerations
Engine mounts, along with other structural components like the motor cases, are designed to withstand the extreme forces and vibrations generated during launch and flight.
Material selection for these components is crucial and includes high-strength options such as steel, aluminum, and alloys like Inconel and Titanium.
3. Comparison with overall vehicle mass
The engine dry mass represents a relatively small fraction of the overall mass of a Delta rocket. For example, the Delta IV Heavy has a total launch mass of 733,400 kg.
Even when considering the mass of the engine mount alongside the engine's dry mass, it would still be a minor portion of the rocket's overall weight.
4. Other types of engine mounts
Smaller-scale model rockets utilize simpler engine mount designs. These use materials like cardboard, phenolic resin, or fiberglass, with associated mass implications.
Important Note: The engine mounts are integral to the rocket's structure, designed to safely contain forces and maintain trajectory during flight. While the precise mass might not be publicly available, the information above helps to understand the scope and scale of this critical component

AI Overview
Detailed public information on the specific mass of the fuel plumbing within the Delta IV rocket family is limited in the provided search results. However, the available information provides insights into the fuel systems and related components, allowing for some inferences.
The Delta IV family of rockets utilized liquid hydrogen and liquid oxygen (LH2/LOX) as propellants for both its first stage (Common Booster Core - CBC) and second stage (Delta Cryogenic Second Stage - DCSS). The fuel plumbing would be a critical part of delivering these propellants from the tanks to the engines.
The search results offer information about the Delta IV's fuel systems and their components:
CBC Construction: The CBC includes isogrid aluminum barrels, spun-formed aluminum domes, machined aluminum tank skirts, and a composite centerbody. A cable tunnel carries electrical and signal lines, along with a feedline for liquid oxygen to the RS-68 engine. This indicates fuel plumbing within the structure.
DCSS Construction: The DCSS also uses isogrid aluminum ring forgings, spun-formed aluminum domes, machined aluminum tank skirts, and a composite intertank truss.
Engine Connections: The RS-68 engine on the CBC connects to the thrust structure via a quadrapod thrust frame and is enclosed in a composite conical thermal shield.
RL10B-2 Engine: The RL10B-2 engine, used in the DCSS, weighs 664 lbs. This weight likely includes internal plumbing but doesn't specify the plumbing mass.
Simplified Design: The RS-68 engine was designed for simplified construction, with a lower chamber pressure, efficiency, and a simpler nozzle than the Space Shuttle Main Engine (SSME). It has fewer parts than the SSME and a simpler nozzle. This might suggest a less complex (and potentially lighter) plumbing system for the engine, but it does not specify the mass of the entire fuel delivery system.
Although the exact mass of the fuel plumbing is unavailable, it is a key component constructed with materials like aluminum alloys and composites. The focus on reduced weight in components like the isogrid tanks suggests an effort to optimize the rocket's overall mass.
Additional points:
The Delta IV Heavy, with three Common Booster Cores, has a larger fuel plumbing system than other Delta IV variants.
In conclusion, although the exact mass of the fuel plumbing is not disclosed, it is an integral part of the Delta IV rocket family's structure and performance, designed with weight and reliability in mind.

As you can see most are not forth coming with all dtails but if you know enough of the dry mass of select parts you can ball park the others.

SpaceNut · Today 14:36:41

AI Overview
Determining the precise gimbal mass for each engine in the Atlas family of rockets is challenging as manufacturers generally don't publicize such specific breakdowns.
Overall engine dry masses can be examined to understand the gimbal systems' contribution to these values:
Atlas V Engines
The RD-180 engine, used on the Atlas V first stage, has a dry mass of approximately 12,000 lbs (about 5,443 kg). It has two thrust chambers that can gimbal independently for pitch, yaw, and roll control, according to Space Exploration Stack Exchange.
The RL10 engine family, used on the Atlas V's Centaur upper stage, has dry masses ranging from 289 to 699 lbs (131 to 317 kg), depending on the specific version. The RL10 also utilizes gimbaling for thrust vector control, with the engines typically able to gimbal up to ±4 degrees.
RL10 Gimbaling System
Early versions of the RL10 engine, such as the RL10A-3-3A, employed a gimbaling system for flight control, allowing for pitch and yaw steering.
Later versions, like the RL10B-2, featured electro-mechanical gimbaling for improved performance and reliability.
In summary, although the specific gimbal masses are not readily available, the engines (RD-180 and RL10) used in the Atlas family are equipped with sophisticated gimbaling systems vital for controlling the rocket's trajectory

AI Overview
Detailed specifications regarding the exact mass of the engine mount structure for the Atlas family of rockets are not readily available in publicly accessible documents.
Some information about related masses is available:
The Atlas I first stage had a booster section with a dry mass of 3,646 kg.
The Atlas II first stage's booster section had a dry mass of 4,187 kg.
The Atlas V Common Core Booster has a dry mass of 20,743 kg (or 21,173 kg for the 55X configuration), which includes the engine and the surrounding structure.
These figures likely include more than just the engine mount itself, encompassing other components of the booster stage structure.
The purpose of the engine mount in the Atlas rockets is crucial for:
Securing the engine during launch and flight.
Aligning the thrust for stable flight.
Providing structural support by connecting the engine to the rocket's airframe and pylon.
Detailed engineering schematics and specific design documents may be needed to find the exact mass of the engine mount structure

AI Overview
Estimating the precise mass of the fuel plumbing specifically within the Atlas family of rocket engines is challenging because this information is not typically broken out as a separate metric in the available documentation.
The overall dry mass of the engines and components can offer insight into the scale involved:
Atlas I Stage 1 (MA-5 propulsion system): Dry mass is 7,882 kg, including a 3,646 kg booster section.
Atlas II Stage 1 (MA-5A): Dry mass is 10,282 kg, with a 4,187 kg booster section.
Centaur Upper Stage: Dry masses range from 1,700 kg to 2,053 kg depending on the variant (e.g., Centaur (Atlas Version), Atlas II Stage 2, Atlas IIA Stage 2).
RD-180 (Atlas III and Atlas V first stage): Dry mass is 5,480 kg.
Atlas V Common Core Booster (CCB): Dry mass is 20,743 kg (or 21,173 kg for the 55X configuration).
Important Considerations:
Engine Plumbing is Integrated: Fuel plumbing is an integral part of the engine and booster assembly, including the turbopumps, fuel lines, valves, and other related components.
Material and Design: Materials like aluminum-lithium skin/stringer and frame for the Atlas III first stage can impact overall weight.
Staged Combustion Cycle: Engines like the RD-180 employ complex plumbing for their high-pressure, staged combustion cycle, which requires an oxidizer-rich preburner and intricate pathways for fuel and oxidizer.
Multiple-Burn Capability: Upper stages like the Centaur with restartable engines have additional plumbing for multiple ignitions and propellant management.
Mass Optimization: Engine designers focus on minimizing the overall mass of the engine systems to maximize the rocket's payload capacity.
Although a definitive number for fuel plumbing mass is not readily available, the figures above provide a general idea of the overall dry mass of the Atlas family's engine components, which includes the plumbing systems.

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