GW Johnson Postings and @Exrocketman1 YouTube videos

tahanson43206 · 2024-04-03 18:35:26

GW Johnson just posted a couple of studies on Single Stage to Orbit and Two Stage to Orbit ....

The spreadsheets can be provided if anyone is interested in seeing them.

Attached are copies in pdf format of two articles I wrote and just recently published on "exrocketman". They are bounding calculations on what can be done toward expendable and reusable vehicles to reach Earth orbit from the surface. One investigates and bounds what can be done SSTO. The other investigates and bounds TSTO. If you want to post these in the drop box thingie, go ahead.
<snip>
This is very simple rocket equation stuff, done in a couple of spreadsheets. The ascent-averaged Isp data come from other stuff I have recently published on "exrocketman", about estimating rocket engine performance. Some of that was done with a version of the bell nozzle rocket spreadsheet used in the orbits+ course. The free-expansion stuff came from a post on "exrocketman" dealing with aerospike nozzles. If anyone wants the spreadsheets, they can have them.
GW

The first file is to be linked from here (SSTO):
https://www.dropbox.com/scl/fi/hu5zc3qc … 3lptv&dl=0

The second file is to be linked from here (TSTO):
https://www.dropbox.com/scl/fi/4dnyyqjn … gsvmd&dl=0

Here is an update:

This is the one that supports the other two you just posted for me. I meant to post this on "exrocketman" some time ago, but never got around to it until today.
It shows vividly just how easy it is to use the "r noz alt" worksheet in the "liquid rockets.xls" spreadsheet file, plus the Paintbrush-made "engine sizing report.png" file to very rapidly size multiple engines and run trade studies. This is where I got my recommendations for how to size a fixed "compromise" bell nozzle to get really good ascent performance out of it.
That spreadsheet file is an update of the one that is part of the orbits+ course materials. I took that one, deleted the extraneous worksheets, added an altitude performance calculation block with automatic plots, and added a output data block that works perfectly with a "Paintbrush"-made engine sizing report. It becomes cut-and-paste with some minor edits to report a design.
The multiple engine designs I sized for the trade study also make a good data library. Pdf document file attached. It should go with the other two, which were "Bounding Calculations for SSTO Concepts" and "Bounding Calculations for TSTO" that you just posted for me.

Link to pdf goes here:
https://www.dropbox.com/scl/fi/s8v6c0zc … upmdo&dl=0

(th)

GW Johnson · 2024-04-08 14:07:02

Despite the clouds, we got to see the eclipse right here in our front yard near McGregor. That's first one I ever actually saw with my own eyes. And I thought I saw a large flare during totality, at bottom right of the sun's image. An orange speck against the corona.

GW

GW Johnson · 2024-04-11 17:36:02

Not sure where to post this. Two news items. (1) the last Delta-4 heavy-lift launcher has flown. (2) The Russians finally on the 3rd attempt got some kind of a heavy-lift launcher to fly.

GW

tahanson43206 · 2024-04-12 21:56:19

We have two new videos arriving soon at exRocketman1 YouTube channel...

I'm in the process of uploading two new MP4 videos to my youtube channel. One of these is the study I did for landing 40-ton cargoes directly on Mars, as 1-way trips. The other is a "how-to" video for using the latest version of the rocket engine performance spreadsheet that was used in the orbits+ course, plus my compressible flow course. This latest version includes the calculations and plots for performance vs altitude, and a sizing of chamber length and diameter (downstream of all the turbopump drive cycle stuff, which is not addressed in the spreadsheet). Cycle needs only the dumped bleed flow fraction to model it effectively in terms of thrust and Isp. These should be visible on the channel after midnight tonight or tomorrow. Youtube has to do its thing processing the MP4's, and each one needs about 3/4 hour to complete. They are about that long.
GW

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tahanson43206 · 2024-04-13 07:23:19

For GW Johnson re new phpBB3 test account ....

I understand that members of the forum do not have time to read every post, so you may not be aware of the test account that the Admins have set up for evaluation of newer forum software.

Please connect to the Azure test account using the information in the Azure topic.

If you encounter an SQL timeout message please just refresh the page. Azure is a Virtual Machine that goes to sleep and has to be re-awakened.

I'd appreciate your feedback on the site.

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tahanson43206 · 2024-04-13 12:10:48

For all...

Here is a link to Dr. Johnson's 42 minute very detailed presentation on how to land 40 (or more) tons on Mars...

https://www.youtube.com/watch?v=lsCZmEiDOB4

At 36 minutes in there is a detailed plan for a mission between Earth and Mars for a reusable transfer stage.

The mission would use liquid hydrogen and liquid oxygen and storable propellant for orbit changes.

The plan assumes a 3+ year flight plan.

At 40 minutes in the summary affirms that 40 tons can be done without a prepared field.

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tahanson43206 · 2024-04-14 07:07:45

Update on latest videos in exRocketman1...

The view count for the long video reported in Post #281 is now up to 1 - that would be my view yesterday

Today I am watching/listening to the video on rocket design ... I would like to point out that since Dr. Johnson learned how to use the on-screen "laser" pointer, he has become increasingly skilled in it's use.

In this video, which is a deep dive into the spreadsheets anyone can use to perform these calculations for design of rocket engines, GW has circled text or cells he wants to emphasize, and he moves the pointer to show where the viewer might focus.

The pace of the presentation in all these videos seems comfortably slow to me, and in any case, the YouTube viewer allows for repetition of sections as desired.

Note for GW Johnson: It occurred to me (while watching the video) that I don't know where to find the spreadsheet, and then up popped a screen showing that the spreadsheet is available at about minute 30 ... I'll wait until the video ends to test the links!

I'm happy to see you are advertising the course offerings, via the links in this slide.

Update a few minutes later... Dr. Johnson! We have a problem! The links do not work from the youTube video.

You might be able to add a page as an appendix to show the links so people can copy them by typing.

For GW ... the links may have pointed to NewMars for folks to pick up materials. It that is the case, please considering including NewMars as a reference.

I'd like to encourage other NewMars members to watch and comment upon these videos.

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tahanson43206 · 2024-04-14 07:40:09

For All regarding links to resources from YouTube video...

Does anyone know how to offer live links from YouTube videos?

Can they be offered via text about the videos?

There may be a convenient way to do this?

For GW Johnson ... there may be a way for you to do this...

No views Apr 12, 2024
This is a quick "how-to" guide for using a spreadsheet file that I developed, which easily and quickly and accurately estimates rocket engine performance. The user does not have to be a rocket engine expert in order to use this spreadsheet. Pretty much anyone could do this.
Transcript

Above is a quote of the description you created for the video... You may be able to put the links in there, or one link to a NewMars post where all the links are stored. You may be able to use this method for all your videos.

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GW Johnson · 2024-04-14 09:41:57

I just did the best I could, with my very limited understanding of all this internet/computer stuff.

The recent posts now show numbers of views. The old one about the Lionel train has 1300 views now, but it's been up there for many years. My son put it there.

For me, it's monkey-see, monkey-do. And the funny part is that I really do like bananas.

GW

tahanson43206 · 2024-04-14 10:46:18

For GW Johnson!

Congratulations on all you've accomplished!

Please see if you can put the links to your spreadsheet in the description field of your video.

I have an offer for you ... if you will put the link to this post into the description field, I will fill it up with all the links you want to provide to your viewers.

You don't have to do ** anything ** other than put the link to ** this ** post into the description field.

The link you want to insert is: https://newmars.com/forums/viewtopic.ph … 34#p221834

You can do this for each video, or for all of them if that makes sense.

NewMars.com has plenty of room for posts so please take advantage of the generosity of Executive Director James Burk and Society President Dr. Zubrin.

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tahanson43206 · 2024-04-14 19:50:23

For All... GW Johnson was successful in registering with the new test forum this evening.

Every current member of the forum is welcome to register with the new test forum.

The address is published at the top of the Azure topic. There are a couple of quirks that everyone should know about ...

Because this is a virtual machine, we are seeing error messages that (I think) are caused by the virtual machine waking up out of a sound sleep, waiting for you to knock on the door. When you ** do ** knock on the door, the Virtual Machine will grumble because it has to load a copy of the MySql server which it only loads when someone needs it. You'll see a complaint that the database missing. Just refresh the page and say Yes to the question if you really want to refresh the page. At that point, it appears grumpy will have had time to load the database.

When you first open the site, you will have to Register. Register is up in the upper right, to the left of log in. The procedure should seem familiar or at least reasonable.

Here is the next quirk... Create a post and you will cause me to see a notification that you have arrived. I will approve your post, and I will also lift you from Newby to Registered member.

At that point, you will be able to help the Admins to evaluate the new system. Please remember that everything we do is temporary, so put your permanent posts into NewMars as usual.

If you have any questions, please post them in the Azure topic.

This topic is supposed to be about GW Johnson's many publications, including the exRocketman1 YouTube videos.

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tahanson43206 · 2024-04-27 18:04:09

This post is reserved for a set of files created by Dr. Johnson in support of the exRocketman YouTube videos, and in support of the course on Basic Orbital Mechanics including rocket design.

There are three files to be linked here:

1) An updated spreadsheet
https://www.dropbox.com/scl/fi/9nqdv47z … f6avf&dl=0

2) A slide show about the spreadsheet
https://www.dropbox.com/scl/fi/uxhf343c … j1suw&dl=0

3)_ A User Manual for the spreadsheet
https://www.dropbox.com/scl/fi/dw1o8end … 0c208&dl=0

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GW Johnson · 2024-04-28 09:08:11

What Tom posted for me in post 287 just above is a revised version of my entry spreadsheet that embodies the 1953-vintage by-hand analysis of H. Julian Allen, used for warhead entry back then. I've had a version of this spreadsheet for some years now, and it is part of the multiple lessons in the "orbits+" course series. What I have done is clean it up of extraneous stuff not used, add a worksheet set up for Venus entry, and I added a model for plasma radiation heating at the stagnation location (which I validated against Apollo experiences).

This is a cleaner-looking, and more complete version. It does radiation heating in addition to convective heating, something not needed in the 1950's with warhead entry. There is the Excel spreadsheet file with worksheets for each of 4 worlds, plus a pdf document that is a user's manual, written to be very detailed and complete. Both of these are better than the original versions. I also created a set of powerpoint slides that is also a sort of user's manual. Those are the 3 things Tom posted for us.

The new spreadsheet file actually has 6 worksheets in it. The 4 "operational" ones for general use are named "Earth", "Mars", "Titan", and "Venus", each being set up with that world's atmosphere model from the Justus and Braun paper on entry, descent, and landing. All you need to analyze an entry are the entry conditions (speed and angle-below-horizontal), and the object characteristics (ballistic coefficient and effective nose radius). You do need to adjust the altitudes in the list (1) to get a denser point spread where things are changing rapidly, and (2) to make the final altitude in the list the one where speed is at local Mach 3, where it is no longer hypersonic (each worksheet has a recommended value listed on it). Plots are generated automatically, but un-annotated.

The other two spreadsheets are copies of the "Earth" model in which I analyzed Apollo coming back from LEO versus Apollo coming back from the moon, in order to "calibrate" the plasma radiation model to match Apollo experiences. I left them in place so all could see how good a match was achieved. But, if one wants to delete them as redundant, that's OK. From LEO, radiation heating was negligible, while it dominates coming back from the moon at just under escape speed. The model predicts just under the nominal 11 gees coming back from the moon, which is just about right.

I would encourage any forums user to download these items and try them out. In particular, the user's manual document shows exactly how I copy the plots to a "Paintbrush" png file and annotate them for clarity. It makes a really good presentation of the results. Earth, Titan, and Venus pretty much come out of hypersonics at high altitudes. It is Mars where you come out low, because of the thin atmosphere. Which in turn makes it very sensitive to ballistic coefficient (higher coefficient is very close to the surface indeed). But all these entries (except Titan) must take place at shallow angles, or else the deceleration gees and heating get too high.

GW

Last edited by GW Johnson (2024-04-28 09:21:15)

tahanson43206 · 2024-04-28 20:02:49

For GW Johnson ....

Per our discussion in the Google Meeting 2024/04/28...

http://newmars.com/forums/viewtopic.php … 67#p222467

Please insert the link above into the description field for the video Mars Direct Landings...

Please provide me with the details I need to insert into the post so that your viewers can easily pull spreadsheets, pdfs and whatever else you want them to have. Please provide text to help your viewers understand what the resources are and how to use them.

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tahanson43206 · 2024-05-02 07:12:22

For All NewMars members who might be interested in the collection of videos that GW Johnson has created and is posting on YouTube.

In an effort to help viewers to gain access to the spreadsheets and other instructional material needed to replicate the rocket design shown in the videos, GW Johnson is considering adding a link to the resources in the Description field of each video.

Here is what the Description field looks like now, for the Mars Lander video:

20 views Apr 13, 2024
This is about how to land large-tonnage payloads on Mars, without any infrastructure there to receive those vehicles. Direct entry off the interplanetary trajectory from low Earth orbit, a low altitude coming out of aerobraking because of the large ballistic coefficient, and a simple rocket-braked landing.
Transcript

Of all goes well, the description field will soon look like this"

20 views Apr 13, 2024
Click here for resources: https://newmars.com/forums/viewtopic.ph … 67#p222467
This is about how to land large-tonnage payloads on Mars, without any infrastructure there to receive those vehicles. Direct entry off the interplanetary trajectory from low Earth orbit, a low altitude coming out of aerobraking because of the large ballistic coefficient, and a simple rocket-braked landing.
Transcript

NewMars can help by hosting the links to resources. This is quite an innovating idea, and to my knowledge, no one else is offering anything like it.

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tahanson43206 · 2024-05-05 07:33:28

For GW Johnson! Congratulations for performing a test launch of a new way to provide information to your viewers!

This first test was not 100% successful.

In many posts of yours about SpaceX you have commented favorably on their willingness to take calculated risks to try out new ideas. You have shown the same quality of leadership in your first test of a new way to deliver resource material to your viewers.

I would describe this as a successful test. You did ** not ** reach orbit, but the booster landed safely back on the Earth, and in looking over the logs of the test, we can see what you were attempting.

As I write this, I do not know the answer, but I think we can make some educated guesses.

The ** first ** observation is that YouTube does not appear to support Hyperlinks in their Description field.

This is disappointing but not surprising, because managing hackers would be a challenge.

The ** next ** observation is that the links your provided have been compressed so they are not usable.

I have no idea (at the moment) where that is happening, but today's Google Meeting will provide an opportunity for us to investigate.

***
For all....

Please take a look at the exRocketman YouTube video description for the Mars Landing video. You will see an innovative design for that field, and you will be able to see what was intended and what worked, as what did not work.

Congratulations to GW Johnson for taking this Bold Leap into the unknown of Internet configuration.

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tahanson43206 · 2024-05-15 11:44:24

For GW Johnson!

The opening text of your post that caused AISE was accepted by the new version of our forum software.

http://40.75.112.55/forums/FluxBB/index.php

Please log into the clone using your regular username and password

After log in, attempt to post the exact same text there.

Please keep in mind that the clone is running data from 2023, so things will look a bit different.

Just pick a topic and post to it.

I would ** really ** like to know if we have solved the AISE problem.

If you post passes muster, then we have added incentive to complete the upgrade.

(th)

tahanson43206 · 2024-05-20 15:19:49

For GW Johnson re reconciliation of two different computations....

Here is a sequence that I found for computation of a volume: =4*$K$5*B11^3/3

Here is another sequence prepared by ChatGPT4o for a Python computation:

5. Volume of a Sphere
The volume of a sphere is calculated using the formula:
Sphere Volume Formula: [ V_s = \frac{4}{3} \pi r^3 ]
V_s: Volume of the sphere (m³)
r: Radius of the sphere (m)
π (pi): Constant value (approximately 3.14)
We now know the volume (V) from the ideal gas law (approximately 3.72 m³). We can use this value to find the radius (r) of the sphere.
Finding the Radius: Set the volume obtained from the ideal gas law equal to the volume of the sphere and solve for the radius: [ V = V_s ] [ 3.72 , \text{m}^3 = \frac{4}{3} \pi r^3 ]
Isolate r by dividing both sides by (4/3)π and then taking the cube root: [ r^3 = \frac{3.72 , \text{m}^3} {\frac{4}{3} \pi} ] [ r = \sqrt[3]{ \frac{3.72 , \text{m}^3} {\frac{4}{3} \pi}} \approx 1.53 , \text{m} ]
6. Conclusion
The volume of a sphere required to hold 1 mole of hydrogen gas at the mean surface pressure (610 Pa) and temperature (0°C) on Mars is approximately 3.72 cubic meters. The corresponding radius of the sphere is approximately 1.53 meters.

What I'm hoping is that my spreadsheet attempt will deliver the same result as the Python attempt.

For a balloon of diameter 1.53*2 (3.06) meters, the volume should be 3.72 cubic meters.

I'd be interested in seeing results that our members might contribute, using a hand calculator or even pencil and paper.

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tahanson43206 · 2024-05-22 18:34:04

For all ... this post is about rocket sizing ... GW Johnson provides a spreadsheet and user manual

I'll put the text after the links...

Here is the spreadsheet:
https://www.dropbox.com/scl/fi/600qk03p … yqmag&dl=0

Here is the user manual:
https://www.dropbox.com/scl/fi/cuvcuddj … iijai&dl=0

Here is an addendum to the user manual:
https://www.dropbox.com/scl/fi/btl8weqm … 0xh7t&dl=0

And here is a pdf comparing SSTO and TSTO designs ...
https://www.dropbox.com/scl/fi/rjv90o1o … pg5ak&dl=0

And here is the text about the links:

(for NewMars members):
As you know, I did a little simplified vehicle sizing spreadsheet “launch sizing.xlsx”, responding to the latest round of SSTO talk on the forums. These were simplified to one weight statement per stage, no linked burn calculations. That rules out complicated missions and reusability. But it works for simple all-expendables on a simple mission: LEO.
I wrote up a user’s manual for it. At that time, there were only worksheets for SSTO and TSTO expendables. The default examples in those worksheets were just approximate guesses. I went back and adjusted those data by running the “r noz alt.xlsx” spreadsheet to get better Isp data, and I revised the TSTO second stage acceleration requirement at its ignition.
Those data got written up in a document titled “Effects of Better SSTO and TSTO Modeling Data”. It covers both LOX-LH2 and LOX-LCH4 SSTO’s, and a LOX-LH2/LOX-RP! TSTO. I was surprised to find the hydrogen SSTO had about the same payload fraction as the TSTO, close enough to be a serious competitor in the expendable launcher market. I was unsurprised that the methane SSTO turned out as poorly as it did. We have known since before WW2 that low Isp requires more stages, and that is exactly the dilemma where the methane SSTO falls.
I then added a third worksheet “tank volumes” to “launch sizing.xlsx” that estimates the tank volumes and stage dimensions for designs created in the “SSTO exp” and “TSTO exp” worksheets. This is designed to work with engine data from “r noz alt.xlsx”, beyond just Isp. The engine configuration is thrust-rescaled to fit the required thrusts coming from the “SSTO exp” and “TSTO exp” worksheets. And the r-value is used to split propellant into oxidizer and fuel masses.
This third worksheet is documented in an addendum document for the user’s manual for the “launch sizing.xlsx” spreadsheet. Attached hereto are 4 of the 5 listed items (there is no need to transmit the original spreadsheet without the tank sizing worksheet):

1. Excel spreadsheet “launch sizing.xlsx” (original: only worksheets “SSTO exp” and “TSTO exp”, not attached hereto)
2. User’s manual: “User’s Manual Launch Sizing.pdf” (cruder examples for “how-to”)

3. Refined models by “r noz alt.xlsx”: “Effects of Better SSTO and TSTO Modeling Data.pdf” (these are still only weight statement results, no tank volumes or stage dimensions)

4. Added tank sizing to “launch sizing.xlsx”: “Addendum to User’s Manual.pdf”, refined examples complete with tank volumes and stage dimensions (this version attached)
5. Revised spreadsheet file “launch sizing.xlsx”: added worksheet “tank sizing”

I suspect you will want to put these things in that dropbox thingie, with links in a posting somewhere on the forums. The text of this letter could go in that posting with those links.
What I found with these simple tools:
When I tried tank sizing for the methane SSTO, I found I needed an engine cluster that simply would not fit behind a stage that met the “aerodynamically-clean” criterion of stage L/D = 6. The much larger takeoff mass (and required thrust) prevents that. It’s not just unattractive from a payload fraction standpoint, it’s actually geometrically infeasible! It would have to be shorter and fatter for the engines to fit, and that is more drag, more drag loss, and a higher required mass ratio-effective dV, which drives payload fraction even closer to zero.
With the advent of TSTO launchers that have partial reusability (SpaceX’s Falcons), which demonstrably impacts cost even more than payload fraction, I see no point to offering SSTO expendable launchers anymore. They are feasible, and competitive in terms of payload fraction with all-expendable TSTO’s, but they cannot offer any reusability, while the TSTO can (flyback first stages), at the “cost” of somewhat-reduced payload fraction. The expendable market will vanish!
And, by the way, when I do the engine ballistics thing versus altitude, I can get a higher ascent-averaged Isp out of an “ascent compromise” fixed bell, than from any conceivable free-expansion nozzle. Meanwhile, bell extensions, which do work, reduce engine thrust/weight considerably. That adds extra inert mass, costing payload fraction out of otherwise the same design. Plus, they add failure modes the fixed bell simply does not have.
GW

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tahanson43206 · 2024-05-26 11:19:05

This post is about McGregor, Texas

GW Johnson lives just outside McGregor, Texas and within earshot of the SpaceX test range.

https://www.yahoo.com/tech/spacex-rapto … 52362.html

by
Mariella Moon

Mariella Moon·Contributing Reporter
Sat, May 25, 2024 at 7:00 AM EDT·2 min read

A SpaceX testing stand at the company's McGregor, Texas facilities went up in flames during a test of its Raptor 2 engines on the afternoon of May 23. According to NASASpaceflight, the engine had an anomaly that caused vapors to seep out and lead to a secondary explosion. The news organization's livestream showed the engine shutting down before the fire started and eventually swallowed the stand in flames and smoke.

SpaceX uses the Raptor engines for its Starship system's Super Heavy booster and upper-stage spacecraft. They use liquid methane and liquid oxygen as fuel, and they were designed to be powerful enough to be able to send Starship to the moon and Mars. As Gizmodo suggests, its gases mixing due to a leak or a similar anomaly could've caused the explosion, though SpaceX has yet to officially address what happened during testing.
The company is currently preparing for Starship's fourth test flight, which is scheduled to take place on June 5, pending regulatory approval and barring poor weather or other factors that could delay the launch. This explosion likely wouldn't affect the flight's launch window. SpaceX's main goals for the fourth test flight are to make sure that the Super Heavy booster gets a soft splashdown in the Gulf of Mexico and to achieve a controlled entry of the Starship spacecraft. The company said it made several hardware and software upgrades to incorporate what it learned from its third flight test. Starship's upper stage reached space during that flight, but it burned up in the atmosphere upon reentry, while its Super Heavy booster broke apart in the final phases of its descent instead of softly splashing down into the ocean.

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GW Johnson · 2024-05-26 12:00:51

Hadn't heard anything I recognized as a failure. Did hear one very short test about 3 days ago, though. That was probably it.

Meanwhile, I have comments relevant to SSTO concepts:

Many correspondents on these forums may make different assumptions than I did, trying to bound what an SSTO can do. It almost doesn't matter, as long as your assumptions are in the ballpark, your result will be, too.

Mission: just assume something like 9.3 km/s of delivered dV covers getting to an orbital destination. That would be for a low-drag configuration with 0.5+ net upward gee at liftoff.

Engine: you probably won't like my modest-technology engines, with only a 2000 psia chamber pressure at the nozzle entrance, only a 2:1 turndown ratio, and about 5% dumped turbopump bleed gas. I've been using an ascent-compromise fixed bell with that, one that is just barely non-separated at sea level and about 80% max chamber pressure. That sacrifices some sea level thrust and Isp for better vacuum thrust and Isp, and pretty much produces a good ascent-average Isp: near 428 s with LOX-LH2. That's a Vex = 4.197 km/s.

You may want to use a bit higher. Fine. Do so. How about 440 sec? Corresponds to Vex = 4.315 km/s. We'll use that.

The mass ratio required to reach destination is exp(dV/Vex) = 8.630. The corresponding propellant mass fraction Wp/Wig = 1 - 1/MR = 0.884, or some 88.4% of the liftoff mass is propellant. THERE IS NO WAY AROUND THAT. My other numbers are right in that same ballpark. Go look for yourself. Go do the calculation for yourself. I just gave you the equations.

That leaves 0.116 (or 11.6%) of the liftoff mass to be both inert mass plus payload mass. THERE IS NO WAY AROUND THAT, EITHER! And my other numbers from the other studies are also right in that same ballpark.

Just remember, the sum of payload plus inert is your dry-tanks burnout mass. Add the propellant to it for your filled-tanks ignition mass. MASS MUST BE CONSERVED! Cannot play games here!

Now if you have 11.6% allowance, and you believe you can build an all-expendable stage with 4.5% inert, which has been done, then that leaves 7.1% max payload fraction. My other studies fell in the 6-8% range, too! Surprise, surprise. LOX-LCH4 did NOT. The Isp is too low, pretty much regardless of what you assume for engine technology. And LOX-RP1 is even a bit lower. Those are SIMPLY NOT candidates for an expendable SSTO application! But LOX-LH2 is, and I got just about the same payload fraction out of a LOX-LH2/LOX-RP1 all-expendable TSTO at 7-8%.

But, based on what SpaceX has already done with its Falcons, the all-expendable launcher market is going to disappear in the next few years. Partial reusability is a bigger contributor to lower cost than payload fraction. We've already seen that in practice.

You have ONLY 11.6% inert+payload allowance using LOX-LH2 in an SSTO, give or take a single % point, depending upon what you assume for your engine and nozzle technologies. SpaceX's "Starship" inert has been running about 120 m.tons in configurations still not fully equipped like a production vehicle. Up to now, the max propellant load has been 1200 tons, and no payload has been carried. That's a stage ignition mass of about 1320 m.tons. The inert is 9.1% at payload 0% and propellant 90.9%. (As I already said, mass must be conserved.)

However, that design has yet to survive entry, and I'm betting its inert mass is going to grow before it is successful. And then they still have to land the thing. So also would a reusable SSTO have to land in some way. That’s going to increase not only the inert but maybe also the dV requirement. Which means your non-propellant allowance is going to go down!

My best guess is that a truly reusable SSTO vehicle that can land by any conceivable means will have an inert mass fraction nearer 15% than any 10% figure. And if that is truly the case, your payload fraction is 2-to-4% NEGATIVE, or worse, which says IN NO UNCERTAIN TERMS that a reusable SSTO is infeasible, even with LOX-LH2.

Even if you could succeed in entering and landing at ~10% inert, your payload fraction will be down under 2%. Partly or fully reusable TSTO's will outcompete you by carrying larger payload fractions than that.

GW

kbd512 · 2024-05-26 15:26:27

GW,

At no point in time has any claim been made that SSTO will beat TSTO on payload performance. That never was and will never be in the cards for SSTO. Basic physics says TSTO is superior for orbital launch. However, that is not the point behind SSTO. The point is to use a singular vehicle to send humans, and only humans, into orbit. For any kind of cargo, we will use TSTO or launch assist technologies (for low value materials such as water, gases, and metals). Human transport is the only plausible use case for spending more money, in order to reduce operational cost and complexity- not the complexity of designing or building a SSTO, but the actual operation of one high-performance rocket powered vehicle vs two, for the express purpose of creating an orbital passenger transport that operates similarly to an airline service.

Falcon 9 is a reusable booster with a 6.06% structural mass fraction for the booster stage. It's primarily Al-2195 alloy, which has a yield of 87ksi. T1200 fiber composite has a tensile / yield strength (basically the same thing) for CFRP, of 572ksi. To that, we're going to add SWCNT to the resin matrix. That bumps up yield / shear / buckling strength by about 75% for a given weight, because the CNT literally locks plies of fibers together by wrapping around them, like almost like a double-ended fish hook (what it looks like under a microscope), to the point that a high pressure CFRP PET-lined O2 tank for a firefighter that previously weighed 57lbs then became 6.1lbs for the same CFRP and PET liner, with the SWCNT resin matrix additive, for equivalent strength. The tank holds back the same pressure as a much thicker ordinary CFRP O2 tank, but it's drastically lighter because it's drastically stronger in every dimension.

If we built the Falcon 9's propellant tanks from T1200 composite, there is no reason I can think of as to why it absolutely must remain at or near the same weight as the Al-2195 propellant tanks. It will be both stiffer and stronger, by quite a lot, as compared to its metal-based counterpart. Al-2195 has 15% of the strength of a T1200 fiber composite, or 17.33% the strength of T1100 fiber composite (current technology for high-strength aerospace parts which are mass produced). T1100 fiber composite (the composite, not the fiber itself, which is 2X stronger than the composite) is both stronger and lighter than any metal alloy that is not more brittle than glass. Weaker but cheaper fibers such as T700, T800, and IM7 have been used by NASA and our primes for reducing weight and total fabrication cost, as compared to Al-2195. PROOF 900HT resin has a maximum wet service temperature of 288C, and is CTE-matched to the fiber, which is far more heat than Al-2195 can tolerate without becoming silly putty, with very little loss of tensile strength up to that temperature (somewhere between 10% and 20% loss of strength, but that only happens during reentry when the vehicle is much lighter, because the TPS provides quite a bit of thermal insulation during ascent). It cures at a much higher temperature than that (454C, IIRC), which is how and why I know that resin won't "go soft" (because you have to get it much hotter to cure it). This has been exhaustively tested by the US military for composite parts in close proximity to very hot jet engines. That is how and why we know that it works. It's been tested- a lot.

UltraMet makes 4.4lb/ft^3 toughened aerogel "Space Shuttle tiles", flown in space aboard the X-37 as part of its TPS package. A 243s exposure to 3,500F / 1,927C for a 1inch thick tile (0.367lbs/ft^2 / 3.3lbs/yd^2), resulted in a backside tile temperature of 350F / 177C at the end of the test. 177C is well below 288C (wet service limit for 900HT) or 350C (dry service limit for 900HT resin). Maximum Space Shuttle reentry temperature was 1,477C. Despite that fact, my mass estimate for TPS included RCC for the leading edges and UltraMet tiles everywhere else, even though lighter flexible fabric reusable surface insulation would suffice. NASA is actively developing BNNT and aerogel fabrics which are absurdly lighter than the flexible reusable fabric insulation used aboard the Space Shuttle.

When the composite strength is 572ksi, you're nowhere near the limit of its strength, your stiffness is more than double that of Al-2195, and the thickness of the composite (5.08mm) is more than double that of Aluminum, which means it's 16X as stiff, as compared to 2195. If Falcon 9's metals-based propellant tank structure doesn't provide that kind of strength, yet it's reusable, then how is a much stronger composite going to fail when its subjected to the same forces during ascent (when the vehicle is heaviest)? What I'm working on is, in point of fact, and despite greatly reduced weight, still far stronger than any equivalent metals-based structure. I've allocated more volume of material that is both stiffer and stronger. How is a Falcon 9 booster reusable when it cannot come anywhere near the strength and stiffness of a composite for equivalent weight? What mechanical properties does Al-2195 imbue the structure with?

My Delta-V with 2,000t of LOX/RP1 propellant, at 316s Isp, is 9,746m/s. The Isp of Oxygen-Rich Staged Combustion RP1 fueled engines, without vacuum nozzles, ranges between 311s and 338s, so I chose an Isp value modestly above sea level, since the vehicle won't remain at sea level for more than a handful of seconds. The majority of its ascent time will be spent substantially above sea level, where the engines produce more thrust per unit mass of propellant expended. I need RP1 engines with demonstrated Isp and a 200:1 thrust-to-weight ratio. That's already been achieved, plus a little extra. I'm not invoking any engine technology which hasn't already existed for quite some time now.

Everything that is not "engine" on the Falcon 9 booster weighs 21,370kg. The landing gear weigh about 2,000kg. By definition, the rest of that 19,370kg has to be propellant tank or thrust structure. It's 7.174 cubic meters of metal if it's primarily Al-2195. If it was made from HexCel IM7 fiber composite, then 7,361kg is a good guess as to how much it would weigh. If it was made from T1200 composite, then 2,906kg (no CNT resin additive) is a good guess as to how much it would weigh for equal strength.

25,600kg dry mass / (25,600kg + 395,700kg) = 6.08% dry mass fraction (as-built by SpaceX)
4,230kg + 7,361kg + 2,000kg = 13,591kg (new dry mass fraction with IM7 composite (as-built by Boeing / NASA)
13,591kg / (13,591kg + 395,700kg) = 3.32% dry mass fraction (IM7 / CYCOM 5320-1 resin)
4,230 + 2,906kg + 2,000kg = 9,136kg (new dry mass fraction with T1200 composite (no CNT additives)
9,136kg / (9,136kg + 395,700kg) = 2.26% dry mass fraction (T1200 composite)

FYI, the IM7 composite, rather than bursting at 1.5X the pressurization load of the Al-2195 tank, burst at 277psi, which means it was absurdly over-built, but was absolutely guaranteed to pass all load tests. Every single one of these composite substitutes for Al-2195 has been overbuilt to the nth degree. Instead of asking why we're over-building composites 2X to 3X stronger than Aluminum (the reason they're not even cheaper), why isn't anybody asking what the logic is behind making the composites ridiculously stronger than any of the metal structures we routinely accept into service?

Boeing used a whole lot of fiber to avoid putting that tank into an autoclave. It could've been much thinner and lighter than it was, using the same IM7 fiber, with improved H2 permeability rates, had they used their giant autoclaves. Yes, it would take longer, no it probably would not be much cheaper than Al-2195 at that point, but performance is what makes the SLS useful, not saving a trivial amount of money relative to the entire project cost.

Anyway...

Would the landing gear still need to weigh 2t for a rocket stage with less than 1/2 of its original dry mass?

The kinetic energy absorbed by the landing gear is the product of mass and acceleration- ye olde F=ma. If the mass is half but vertical acceleration rate is the same, then you have half the kinetic energy. The vehicle is also dramatically "butt heavier", thanks to the engines and thrust structure now comprising almost 46% of the entire dry vehicle weight, so track width can be narrower for equivalent stability, which further reduces weight. We'll ignore that optimization and focus solely on weight added to make the gear strong enough.

8,136kg (gear now half as heavy, absorbing half as much energy on landing) / (8,136kg + 395,700kg) = 2.01% dry mass fraction

Would we be modestly "heavier than that" after accounting for our engine thrust structure and grid fins?

Yes, but not by a lot. That means a 2.75% dry mass fraction is achievable using 1,000ksi+ composites (T1200 fiber plus SWCNT resin matrix additive) and 200:1 thrust-to-weight ratio RP1 engines with a non-regeneratively cooled RCC nozzles plasma sprayed or CVD coated with UHTC, to prevent oxidation, as already proven to work and reduce weight and engine complexity by NASA engine testing with these nozzles.

The notional SSTO I have in mind is 4.31% payload and dry mass fraction. The 2,000t of propellant provides 9,746m/s of ΔV at 316s fixed Isp (reality is that Isp improves by a little bit more during ascent, up to 338s, 100bar chamber pressure). No aerospike, no nozzle extensions, etc. That should be enough to make up for drag and gravity losses.

To reiterate, I have completely conceded the point that a TSTO will outperform a SSTO, every single time. That is not "the why" behind building a SSTO. There are other factors at play beyond simple fuel cost and payload performance.

Can we also apply our same SSTO composite tech to TSTOs?

We obviously can, and we already have companies like Rocket Labs doing that with their LOX/RP1 TSTOs. We'd be foolish not to, assuming performance matters so much that we're de-justifying SSTOs on the basis of payload performance per unit of dry vehicle mass. It can't possibly be the cost of the composite tech, because we're using that more and more over time, regardless of SSTO vs TSTO. For me, this is more about advancing rocket tech than whether or not a SSTO can deliver more payload than a TSTO. The answer to that question is beyond obvious.

I want to roll our best composite tech, best engine tech, best thermal protection system tech, and best avionics / life support / power / thermal management tech into a single vehicle, to show the entire world what "the best" actually looks like and how astonishing the end result is- something previously thought impossible actually is possible with modern materials. Whether or not it's practical is a different question.

tahanson43206 · 2024-05-26 20:49:17

This post is reserved for a planned update with links to studies relating to SSTO...

https://www.dropbox.com/scl/fi/bniu2fn1 … kj1bl&dl=0
A Quick Look at Reusable SSTO

https://www.dropbox.com/scl/fi/nxefg4sa … nj4yk&dl=0
Addendum-2

https://www.dropbox.com/scl/fi/kion5sgd … 0gpvt&dl=0
Addendum to User Manual

https://www.dropbox.com/scl/fi/ezudh1fr … yjz9u&dl=0
Descriptive text SSTO studies

https://www.dropbox.com/scl/fi/xdk29356 … amin2&dl=0
PDF version of SSTO text

https://www.dropbox.com/scl/fi/ufyre58d … 99zms&dl=0
The spreadsheet to make it all happen ... Launch Sizing

https://www.dropbox.com/scl/fi/tfquke9u … 0t8y5&dl=0
User Manual for Launch Sizing spreadsheet

https://www.dropbox.com/scl/fi/ndlswwdn … 0thf2&dl=0
What the Bounding Calculation for SSTO Ultimately is

(th)

kbd512 · 2024-05-27 02:07:36

"Densified" (Oxidizer and Fuel Melting Points + 10C) Bulk Propellant Densities with Proper Mixture Ratios
LOX: 1,262kg/m^3; RP1: 867kg/m^3; LOX/RP1: 1,124kg/m^3
LOX: 1,262kg/m^3; LCH4: 438kg/m^3; LOX/LCH4: 890kg/m^3
LOX: 1,262kg/m^3; LH2: 77kg/m^3; LOX/LH2: 395kg/m^3

Staged Combustion Sea Level Isp, Vacuum Isp, Straight Average Isp for Fuel of Choice
LOX/RP1 (RD-180): 311s; 338s; 324.5s
LOX/LCH4 (Raptor 3): 327s; 356s; 341.5s
LOX/LH2 (RS-25E): 366s; 452s; 409s

Propellant Mass and ΔV Results with 50,000kg Vehicle Mass and 1,000m^3 of Propellants of Choice, using Straight Avg Isp
LOX/RP1 Propellant Mass - 1,124t; Vehicle Mass - 50t; Total Mass - 1,174t; ΔV Capability - 10,044m/s
LOX/LCH4 Propellant Mass - 890t; Vehicle Mass - 50t; Total Mass - 940t; ΔV Capability - 9,825m/s
LOX/LH2 Propellant Mass - 395t; Vehicle Mass - 50t; Total Mass - 445t; ΔV Capability - 8,768m/s

Demonstrated Engine Sea Level Thrust-to-Weight Ratios
RP1 (Merlin-1D): 184.5:1
LCH4 (Raptor 2): 143.8:1 (Raptor 3 has achieved 169.4:1 in testing)
LH2 (RS-25E): 53.4:1

Engine Percentage of Vehicle Dry Mass Fraction to Achieve 1.5:1 Thrust-to-Weight at Sea Level
RP1: Thrust Required - 1,761t; Engine Mass (184.5:1) - 9.545t; Percentage of Vehicle Mass: 19.09%
LCH4: Thrust Required - 1,410t; Engine Mass (169.4:1) - 8.323t; Percentage of Vehicle Mass: 16.646%
LH2: Thrust Required - 667.5t; Engine Mass (53.4:1) - 12.5t; Percentage of Vehicle Mass: 25%

RS-68A, RS-83, J-2X, Vulcain 2.1, and RD-0120 all have inferior thrust-to-weight performance when compared to the RS-25. There may be a few experimental LH2 engines I don't have numbers for, but I doubt they were any more performant than our RS-25, else they'd have been touted for their performance, far and wide.

Why does thrust-to-weight matter so much for SSTO?

LH2 is 1/2 the density of the lightest hydrocarbon fuel, and all engines burning LH2 are 3X the weight of any hydrocarbon burning engine, because the pumps required to feed the engines are massive relative to pumps feeding much denser fuels. LH2's 114s Isp advantage over RP1 simply cannot make up for the heavy engines and heavy fuel tanks required to provide equivalent payload performance. Greater Isp is an enabler for SSTO, but heavy engines and excessively voluminous fuel tanks are not, because both of those problems eat so heavily into the available vehicle dry mass fraction, and therefore payload fraction. It's counter-intuitive, or at least it was for me.

Using LH2 is stacking the deck heavily against SSTO, as well as TSTO if LH2 is used in the first stage. LH2 with 70% CH4 gellant by weight, only gives up 50s of Isp to neat LH2, meaning 410s vs 460s, yet it's density (177kg/m^3 vs 77kg/m^3) and O/F ratio (4:1) is such that it out-performs neat LH2. Your tank volume ends up between LH2 and LCH4 for equal ΔV, but you have much more propellant mass to work with. I would imagine that thrust improves. Presumably, you'd have much lighter turbopumps and therefore engines because they're sucking down a more viscous gel instead of a rarefied liquid with almost no mass per unit volume. That means you can have lower tank pressures, too, which helps reduce tank mass, since tank pressurization is one of the major loads to contend with.

O = Oxidizer
F = Fuel
MR = Mixture Ratio
ρ avg = (ρO * ρF * (1 + MR)) / ((ρF * MR) + ρO)

I think that's the correct formula to use to determine propellant bulk density. Gelled LH2's bulk density is about 567kg/m^3 if I did my math right. I get the right numbers for LOX/LH2 at 6:1, so I presume it works. My straight average Isp is 363s, and my ΔV capability is therefore 8,945m/s with a 50t vehicle mass and 1,000m^3 of propellant.

Relative to RP1, using LCH4, I need 7% more propellant tank volume for equal ΔV, but only 78% of RP1's propellant mass. There might be a worthwhile trade here, presuming near-identical engine thrust-to-weight, which is presently 91.8% that of RP1, at sea level. Careful evaluation of what engine dry mass does to total vehicle mass is still required. It's probably not possible to make the engine as performant as a RP1 engine, but if it's close enough then that might be good enough.
Relative to RP1, using 70% gelled LH2, I need 39% more propellant tank volume for equal ΔV.
Relative to RP1, using neat LH2, I need 42% more propellant tank volume for equal ΔV.

Needless to say, increasing tank volume by 40% is disastrous to the vehicle's dry mass, because stiffness becomes the limiting factor, rather than strength. Modern composites provide strength galore. Their ability to resist deformation, an intrinsic property of the fiber and resin matrix, is only on-par with high strength steel (not by unit weight, obviously). The low thrust-to-weight ratio of LH2 engines only adds insult to injury.

My propellant load is only 561.7t for LH2 vs 1,224t for RP1 for equal ΔV, but increasing volume by 40% is far more problematic than the strength of CFRP. Isp is clearly much better with LH2, so that's good, but even if I use 90% of vacuum Isp for LH2, I can't do anything about the tank volume. At best, my tanks are 38% larger than they are with RP1, using 90% of vacuum Isp for LH2. I know that cube-square works in my favor with LH2, but is that enough to overcome a 40% volume increase? That seems unlikely. My thrust is also 1/4 to 1/3 per unit of engine weight, as compared to RP1, which would be fine if both the tanks and propellant were lighter to compensate, but LH2 propellant weight is only 1/2 of RP1 for equal ΔV, ignoring all other factors. Basically, LH2 is no help to me at all. I can't overcome the volume and thrust-to-weight problems with higher Isp. Using RP1, I can devote more volume of material to enclose a sufficient volume of propellant capable of producing the required ΔV, because the propellant is 3.1X denser and the mass of my engines 3.46X lighter for sufficient thrust.

All those compounding factors make LH2 very unattractive for SSTO. On top of that, there's a long term issue with LH2 permeation of composites. RP1 doesn't have that problem. LOX isn't even much of a problem. LCH4 has free Hydrogen molecules within it, but not that much. For a TSTO upper stage, LH2 is much more attractive, specifically because the booster has to accelerate its own weight plus the upper stage weight. That's where LH2 really shines, because raw thrust is not what produces spectacular upper stage payload performance. About 80% of a TSTO's liftoff mass is the booster stage, so throwing a lightweight upper stage on top, with stellar Isp, really makes it perform quite well for high orbits and escape trajectories. It's a different class of problem, though. You can't fix the booster thrust and propellant density problem with Isp. Incidentally, nuclear engines suffer from the same problems in booster applications- their thrust is too low to make up for their weight. Gravity losses dominate. LH2 nuclear SSTO would be even worse than a conventional LOX/LH2 SSTO.

GW Johnson · 2024-05-27 10:35:29

I'm old and obsolete, but in my days there were no materials that even came close to 572 ksi tensile strength, and there were no organic matrix composites that could withstand steady service temperatures above 300 F. The highest strength steels back then were in the 280 ksi class, and all the then-known composites no better than mild steel. The best organic bound material we had available was asbestos-epoxy, and it became junk at 290 F. You had to trowel it on thick, to meet burn time (2-10 sec) as a protectant inside rocket motors.

Let's just say I am an open-minded skeptic about such material claims, especially given the way corporations out-and-out lie in their advertising hype. They will say literally anything to make sales, especially to the government. I saw that happen for decades in the defense industry.

The point of my bounding calculations was to show that getting 9.3 km/s out of a single stage was possible with Isp in the low to mid 400 sec range, but not possible in the low-to-mid 300 sec range. It is possible in the high 300-sec range, but you have very little inert+payload allowance left over. Too low to be in the least credible. Not when the indicated propellant mass fraction is 88+ %, and above 90% as the Isp drops. 100 minus that propellant percentage is all the allowance you have! Your inert and payload fractions have to fit within that. And bear in mind that something surviving one entry is still a hell of a long way from providing a long useful service life!

Entry isn't a single temperature at this-or-that location. It is literally a heat balance among some 5 items, not all of which are present in any given scenario. Consider this: at around Mach 10-to-12 around 50 km up, most shallow-angle entry vehicles are close to, or in, max heating. Depending upon shape and ballistic coefficient, that convective heating rate can be anywhere in the range of 300-3000 W/sq.cm.

It's a short pulse, only several seconds long, but it is several seconds long! It will not be equilibrium, but an equilibrium model is a decent guide to what you have to fight. And a rough guide to the effective driving temperature of the plasma in the sheath about that vehicle is in the range of 3300-4000 K.

If that balance is not achieved, the material is going to soak out locally to that driving temperature, much faster than conduction or convection inward to a heat sink can possibly occur. Your only real, demonstrated, cooling options are high-temperature re-radiation or fast ablation, right from the heated surface. Transpiration cooling analyzes as possibly a good third candidate, but has never actually yet been flown to verify that claim.

I'm not saying that an SSTO few-people-only mover to LEO is impossible. I'm just saying it won't be very attractive, and the design space is very limited indeed. You'd be better off with a TSTO and a few people in a simple capsule. Something we already know works. With a wide range of design options available.

GW

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