GW Johnson Postings and @Exrocketman1 YouTube videos

tahanson43206 · 2023-10-01 12:10:20

For GW Johnson re #175

Congratulations on your work on the heat transfer course!

***
The Space Shuttle Main Engine is an example of a design that pushed water in the form of steam to produce thrust.

The steam was created through a chemical process, which is something you know well, and can teach to others.

I thought it might be helpful to ask ChatGPT for key facts about the operation of the Space Shuttle Main Engine.

The chat at the link below collects key facts, including:

Temperature of the steam at the nozzle
Velocity of the steam at the nozzle
Mass flow of the steam at the nozzle.

http://newmars.com/forums/viewtopic.php … 11#p214311

Please review the transcript and confirm if you agree with the figures quoted.

The challenge we have here is to make a transition in thinking that may not be possible for humans to make.

ChatGPT is not human, so it can ** think ** of things that humans cannot.

(th)

tahanson43206 · 2023-10-01 12:25:57

For GW Johnson and all who may be interested ...

Here is an early report on the SlingLaunch system, before the most recent full scale test:

https://www.universetoday.com/153342/sp … -to-orbit/

The system is a one-third-scale model of the OLS system that is currently under development and relies on the same components. Like the OLS, the Suborbital Accelerator uses a vacuum-sealed centrifuge to spin a rocket and then catapult it to space at up to 8,000 km/h (5,000 mph). The rotational kinetic energy comes from ground-based electricity provided by solar and wind (which would eliminate the carbon footprint of rocket launches).

With this first test under its belt, SpinLaunch will be moving ahead with the development of its full-scale accelerator, which will measure 100 meters (328 ft) in diameter and be capable of launching payloads in the 20 to 200 kg range (44 to 440 lbs). The type of payloads they envision includes various kinds of satellites, space-based solar arrays, and electric propulsion modules.

The SlingLaunch company is demostrating the potential to impart velocity and direction (vector) to a mass of several kilograms, using a rotating energy accumulator to achieve a velocity of (on the order of) 5,000 miles per hour.

Meanwhile, medical and commercial centrifuges are routinely operating at 100,000 all the way to 1,000,000 RPM.

The arm length of the commercial centrifuges is less that that of the SlingLaunch system, and the mass in rotation is less, but the important fact that (I hope) will come across is that the velocity at the tip of one of the smaller machines is the critical factor in ISP.

Thrust is a function of the mass thrown in combination with the ISP.

I read recently that the ISP of a photon (as projected from a flashlight for example) is measured in years.

The mass of the photon is so small that it literally cannot be measured except by the most exotic of instruments.

I have requested information about the performance of their high performance centrifuges from a manufacturer in Germany, and will immediately report any response I receive.

(th)

tahanson43206 · 2023-10-01 20:09:59

For GW Johnson ... thank you for the insight you provided by email. In your review of ChatGPT's "idea" you recognized that the "idea" is a member of a set of methods called "Impulse" propulsion. this class of methods is distinguished from continuous propulsion, such as provided by a rocket engine burning oxygen and hydrogen over a period of time. In the transcript at the link below, ChatGPT (3.5 this time) describes how the area under the curve of the respective propulsion methods is what sums up to provide the space vessel with change of momentum. If the area under the curve of the discontinuous method is the same as the area under the curve of the continuous method, then the total momentum change will be the same.

https://www.dropbox.com/scl/fi/ztmtk8bu … l5vkq&dl=0

(th)

tahanson43206 · 2023-10-04 07:25:25

For GW Johnson --- in recognition of your objection to the idea of "impulse propulsion" as contrasted to "continuous propulsion" I have undertaken a search for a potential supplier of centrifugal water pumps able to deliver continuous flows of water suitable for space propulsion. I am prepared to contact the available suppliers, to see if any of them are willing to make the investment that will be needed to deliver space capable pumps.

Please prepare specification of water pumps you would like to have available to propel the Large Ship of RobertDyck, and your Space Tug, which will be capable of accelerating a 5000 metric ton space craft to near escape velocity.

Please put your specifications in the new Centrifugal Pumps topic.

An example of what I'm looking for is a specification for a device that can deliver one kilogram of water per second at 5000 miles per hour.

Your specification will be similar, but tailored to your requirements for Large Ship and Space Tug.

I am hoping your specifications will tend toward large numbers of small pumps, instead of a very few large ones.

We already know (or at least have reports) that centrifuge equipment for medical, commercial and military use can reach 1,000,000 RPM.

The limits of what can be achieved are set by the properties of matter, and by the ingenuity of human beings.

(th)

GW Johnson · 2023-10-04 08:57:55

Tom:

I don't "object" to impulse propulsion. I'm actually a fan of nuclear explosion propulsion, which is quite the extreme example of it.

Like the explosion drive, for your centrifuge sling notion, the proper time interval over which to "average" the pulses of impulse (and the effective rate of thrown mass) is the time between impulses.

With the nuclear explosion drive, (and your centrifuge sling) there is a peaky pulse of applied force as the pulse of applied impulse occurs. With the nuke drive, there had to be a huge two-stage shock absorber between the main ship and its blast-receiving pusher plate. By designing the time constant for shock absorption and motion recovery to be the same as the pulse interval, they were able to smoothe the vehicle acceleration trace to something survivable. As I recall, the 5000-10,000 ton designs had ship acceleration averages in the 2 to 4 gee range, and Isp's in the 5000-10,000 sec range. It only works good at high ship masses, they found. At 500-1000 ton for a NASA Mars Mission, it was only about the same as NERVA in terms of Isp.

Your centrifuge slinger will have a similar pulsed acceleration trace without some sort of shock absorber between where the slinger is mounted, and the rest of the ship. There's a sharp peak with nothing between peaks. The idea is very similar in its effects to what General Atomics's crew was dealing with on old Project Orion (the nuclear explosion drive), except for you it's not so unsurvivably extreme. Your numbers are much smaller.

For continuous ejection of water by pumps, the mass ejection speed you can achieve is limited by the pressure at the outlet of the pump, and the Bernoulli equation, to that which occurs for proper acceleration to the ambient pressure level, even if it is zero in vacuum. It will be far, far short of your 5000 mph figure. That would require mega-pressure, and there are no high flowrate pump designs that produce it. A 1-stage centrifugal design has a max pressure ratio of about 4.

Edit update a few minutes later: the operative equation to relate Voutlet from the nozzle (the massflow ejection speed) to pump capacity is from Bernoulli as dPpump = 0.5 density(Vnozzle^2 - {mdotpump/(density*Apumpoutlet)}^2), where the pump capacity is defined as its pressure rise dPpump at pumped massflow mdotpump, with a pump outlet area Apumpoutlet. This gets fed to a convergent nozzle that reduces the pumped stream conditions to the ambient atmospheric pressure, which could be zero. The pump pressure rise is dPpump = Ppumpoutlet-Pambient. And THAT is what sets the nozzle exit speed Voutlet, for a given set of pump capacity data. The thrust of that nozzle will be the exit momentum F = mdotpump*Voutlet.

About the largest practical water pressures from pumps (multistage centrifugal) that deliver high flow rate, are near only 175 psi, and that's typical of fire-fighting equipment. There's higher pressures, to be sure, but NOT at high flow rates, which you also need, to produce thrust.

Now-defunct XCOR Aerospace was using reciprocating piston pumps for one of its engine designs, but ran into a size limit for the engine thrust that was quite low. This was the flow rate limitation at high outlet pressures for that kind of positive displacement pump. As you scale up engine size, the pump grew faster than the engine did.

GW

Last edited by GW Johnson (2023-10-04 09:48:18)

tahanson43206 · 2023-10-27 09:20:57

For Dr. Johnson!

Please take a look at this conversation with BARD....

It contains errors, which I am hoping you will enjoy finding. The estimate provided by BARD for the apogee of a payload launch from a ballistic launcher able to reach 300 km at perigee is 45,600 km, if the objective is to minimize impulse needed to raise the perigee to 300 km.

If you decide to take this on, I'd be interesting in knowing where the perigee would be if no impulse is added at apogee. I'm guessing perigee is inside the Earth, but I'm not sure.

(th)

tahanson43206 · 2023-10-29 10:08:43

The link below points to a paper that shows study of "parking" orbits for rendezvous ...

Accumulated Orbit Data from the Spreadsheet pdf

https://www.dropbox.com/scl/fi/djuzf7ei … jbl8s&dl=0

(th)

tahanson43206 · 2023-10-29 20:35:37

For GW Johnson re rendezvous on orbit ....

In this evening's Google Meeting, you showed us work you are doing on identifying optimum flight paths for delivery of payloads (propellant) to a station at 300 km. In your example, you showed us 140 km as being above the atmosphere.

Please add to your list of scenarios the present altitude reached by the existing SpinLaunch equipment. This is NOT the final version of SpinLaunch, but it exists today, which means it could deliver payload to the refueling station, even if the quantity shipped is modest.

The altitude reached so far is 40 miles or 64 km. The amount of mass that must be allocated to circularizing the orbit will be much greater than 10%, but I'd be curious to know if it is even possible.

As I envision the burns, there would be two ... the first would create an ellipse with perigee at 64 km, and apogee at 300 km. The second would burn at 300 km and it would raise the perigee to 300 km.

I don't know the mass of the existing test projectiles but let's assume 100 km to work with. Given that mass, and the ability to reach 64 km, can you work out a solution that delivers 1 kg of payload to the station? The mass of the carrier would be bonus, for whatever value that might have. It would be high quality metal, for sure.

It might even be worth collecting the carriers and returning a bunch of them to the launch site for re-use.

(th)

tahanson43206 · 2023-10-30 18:23:52

GW Johnson has been working on rendezvous orbits for delivery of propellant to a refueling station from ballistic launchers.

Below is a link to a pdf file that contains text and two images, showing "inside" and "outside" docking techniques. I'm planning to provide links to the images separately.
https://www.dropbox.com/scl/fi/h5ltdu0q … 1ixwg&dl=0

Image of Inside rendezvous to be added:

Image of Outside rendezvous to be added:

(th)

GW Johnson · 2023-10-31 08:43:38

In post 184 just above are two images I created that Tahanson43206 has posted somehow. They relate to two strategies for approaching a station in LEO with a payload launched from the surface. It does not matter how it is launched, as long as the end of the launch trajectory is the required perigee velocity at an altitude of 140 km (the entry interface altitude), and a path angle of 0 degrees above horizontal.

The "inside orbit" case in the first figure shows a transfer ellipse that is also the rendezvous "chase" orbit if you miss the tight launch window (at max 7 minutes long) to arrive at the station just as it is there, too. It grazes the atmosphere, and is subject to orbital decay and early entry, but you can make several trips around this orbit before losing significant apogee altitude. It has the lower as-launched speed requirement, the lower dV to circularize at the station, and the longer "chase" time if you miss the launch window and have to do the rendezvous "chase" thing. Presumed here is low-inclination eastward orbital direction, because I used Earth's equatorial radius, not its average radius.

The "outside orbits" case in the second figure uses a transfer ellipse that grazes the atmosphere (at the interface altitude) at its perigee, with an apogee at the nominal base of the Van Allen radiation belts. This apogee limitation is for payloads sensitive to high radiation exposure, be they manned or unmanned. It is very closely related to the actual rendezvous "chase" orbit, an ellipse between LEO and and the nominal base of the Van Allen belts. There is a very small apogee burn to adjust the perigee from 140 km to 300 km (nominal LEO). It has the higher as-launched speed requirement for getting onto the transfer ellipse at the edge of the atmosphere, and the higher dV value for a perigee burn to circularize from the "chase" ellipse onto LEO circular, when the station is there. But, it has a significantly-lower rendezvous "chase" time if you miss the launch window.

I ran these orbits with the spreadsheet tool that is part of the "orbit basics +" class that I put together. That spreadsheet just semi-automates the hand calculations once used to calculate such things, using the textbook equations for 2-body orbital mechanics. It (and the course materials) are available here on the forums. Those are located in the Acheron Labs section, "interplanetary transportation" topic, "orbital mechanics class traditional" thread. It's in some sort of drop box thing on the internet. Tahanson43206 put them there for me; I do not understand any of that sort of thing.

It does not matter how you launch the payload to use these orbits to reach the station, just as long as your end-of-launch trajectory values match up with the perigee of the transfer orbit you are going to use. Could be a rocket launch, a light gas gun, or a centrifugal sling device.

For the case of a properly-designed rocket launcher, the mass ratio-effective dV for "inside orbits" case is about 8.8 km/s (and the "outside orbits" case higher still at almost 9.2 km/s), using the entirely-empirical estimating techniques that I also cover in the class materials. I personally do NOT know any empirical estimating techniques for light gas gun or centrifugal spin launch methods! I only know the empirical stuff for rocket launch.

The accurate way to do this launch thing is a computer trajectory code that has at least 3 degrees of freedom, is written in polar coordinates, and has models for thrust vs time and vs altitude, aerodynamics vs Mach number, mass properties vs propellant burn-off, and the physical dimensions that support the rest. I do NOT have one of these computer codes!

But if you have one of these codes, it would also cover light gas gun launch or centrifugal sling launch, given the right initial velocity and no thrust, but with all those other inputs appropriately evaluated. Building the appropriate collection of modeling inputs is a lot of work, I caution you! Much or all of these inputs change, every time you change what it is that you want to "fly" in the code. I know, I used to do this kind of work decades ago.

I could do the same thing looking at transfer orbits that I did here, for low Mars orbit (or for low lunar orbit). But if you use the class materials and spreadsheet, you can do this orbits stuff for yourself! That's what the class materials are for, so you can do this stuff for yourself! The class materials are ultimately aimed at determining ideal dV values from orbit speeds, and then appropriately empirically factoring those dV values up, to cover gravity and drag losses for rocket flight.

Those factored-up dV values are the mass ratio-effective dV values needed for using the rocket equation (correctly) to size (along with realistic thrust requirements) your launch vehicles. There's a lesson in the course specifically about that.

There's also a lesson in the course (and the spreadsheet, too) for correctly estimating rocket engine Isp. The published tables vs propellant combination always make assumptions about nozzle performance that do not match what the user intends. They are therefore nearly always wrong by a few percent, unless you know how to estimate the right values. The lesson tells you exactly how to do that job right, gives you the spreadsheet with which to do it easily, and has a library of propellant combination data.

Update 11-1-2023: I forgot to mention the effects of turbopump drive cycle as an error in the published Isp tables. Very few engines are full-flow with no bleed dumped overboard, even today. The tables presume no bleed dumped. That's a percent off the Isp for every percent of flow from tankage that ends up dumped overboard.

There is also the error of chamber pressure being different in the tables from what you intend. By "chamber pressure" I mean the total pressure approaching the entrance to the nozzle contraction. There are two effects leading to errors in Isp: (1) a direct effect on the nozzle thrust coefficient, which is larger for higher pressures because the expansion ratio can be larger, before you bump into whatever constrains expansion, and (2) the chamber c* characteristic velocity is higher at higher pressures, because it correlates as a power function of chamber pressure. The effect is weak, but real, and can be significant (like 2% just for the c* effect). This is EXACTLY why SpaceX went to 4400 psia in their Raptor-2 designs. Most of the tables are for 1000 psia at sea level and 100 psia in vacuum. End update 11-1-2023

GW

Last edited by GW Johnson (2023-11-01 10:59:44)

tahanson43206 · 2023-11-07 11:10:58

For GW Johnson re question of finding individuals who want to study (and are capable of studying) your course material.

I thought that your use of simple, common, ordinary spreadsheets to perform complex calculations would make the course material accessible to members of this forum, and hopefully to others outside the immediate membership.

Clearly we were mistaken. I just heard from the gent at North Houston NSS who put us in touch with a high school student who seemed to have an aptitude and an interest in the subject matter, but that person did not follow through.

We are left with what I believe is a well-written introductory course on Basic Orbital Mechanics with no students.

I can understand the members not following up on this opportunity. Persons in the later stages of their lives do not want to learn anything new (as a rule - there are exceptions). We have no younger persons in the membership.

Next Saturday, if you tune into the NSS chapter meeting via Zoom, I'd like for you to think about the venue as a possible avenue for reaching prospective students. Please note that everyone in the room, and everyone in the Zoom community has their own agenda. No one is thinking about helping you, with the sole exception of me.

The question before us is the eternal question facing all marketing personnel ... how to reach the customers who want our product, at a price we can afford.

There may be an opportunity if one of the chapter people has an agenda upon which you can hang your offering. As it happens, at least one member is interested in reaching out to young people, but my impression is that this is a one man band operation. No one else is invited to assist. There might be a way to show how your offering could accrue to the glory of this individual, but it will take some convincing.

(th)

tahanson43206 · 2023-11-10 15:12:03

For All ....

The North Houston Chapter of the National Space Society has decided to allow one of their members to attempt to create a series of educational videos based upon the work of GW Johnson.

It has been decided to use Google Meeting as the venue for this work.

To the best of my knowledge, no one involved has attempted something like this, beyond the recorded presentations for Mars Society that GW has already done.

The unique feature (as I understand the plan) is to feature 5 minute knowledge bursts based upon slides an narration, followed by live interactive work sessions in which attendees attempt to make the required mental leaps needed to understand GW's material.

(th)

kbd512 · 2023-11-12 20:19:13

GW,

I'll add documents to this post, related to LOX / LCO (Liquid Carbon Monoxide) rocket engines, as I collect them:

Carbon Monoxide and Oxygen Combustion Experiments: A Demonstration of Mars In Situ Propellants, by Diane L. Linne, Lewis Research Center, Cleveland, Ohio

Experimental Evaluation of the Ignition Process of Carbon Monoxide and Oxygen in a Rocket Engine, by Diane L. Linne, Lewis Research Center, Cleveland, Ohio

kbd512 · 2023-11-12 22:52:50

GW,

Hold off on doing any work on LCO as a propellant until I can recompute my back-of-the-envelope calculations to confirm what ballpark we're playing in, if we opt to use LCO as our propellant. My 1,250kg/m^3 figure was way off, but the figures I keep seeing people, including engineers, use for the bulk density of LNG or LCH4, are equally way off.

Sub-cooled propellant bulk densities (any colder and it freezes):
LOX: 1,306kg/m^3 at -219C
LNG (CH4): 448kg/m^3 at -180C
LPG (C3H8): 733kg/m^3 at -188C
LCO: 849.5kg/m^3 at -205C (NOT 1,250kg/m^3; this was my fault for not checking my source against two other sources)

I've seen LNG bulk density conflated to be 657kg/m^3 (only achievable with LPG) all over the place. Petroleum Engineering companies put the bulk density of LNG between 430kg/m^3 and 480kg/m^3, due to various impurities (Ethane, Propane, Butane, etc).

GW Johnson · 2023-11-12 23:36:18

Kbd512:

I downloaded your 2 documents. Thanks.

I've been using densities from my ancient P&W Handbook from 1969. Those are not subcooled, such was unknown back then. LOX is listed as sg = 1.143, and LCH4 as sg = 0.424. Those would translate to 1143 kfg/cu.m and 424 kg/cu.m. I used those as "conservative" in the sense of not pushing the state-of-the-art.

Will hold off on looking at LCO.

GW

Last edited by GW Johnson (2023-11-12 23:39:14)

GW Johnson · 2023-11-13 09:15:01

I looked in the first document of the rocket performance tests (second concerned ignition). I was able to correlate c* vs Pc, but not r. At 1000 psia, c* is near 4150-4160 ft/sec for LOX-LCO, vs about 5900 for LOX-RP1, 6120 for LOX-LCH4, and about 7950 for LOX-LH2. All of these are weak power functions of Pc with exponents near 0.01. But let's ignore that effect for now.

This figures into Isp along with the nozzle performance as Isp = [c* CF (1-BF)]/[CD gc] when you use the real-world effects. For a well-designed nozzle CD ~ 0.995, and gc is just the units conversion constant for the inconsistent units in which Isp is defined, whether metric or US customary. BF is the fraction of propellant flowrate drawn from tanks that ends up dumped overboard instead of going through the nozzle. In a full-flow cycle, BF = 0.

CF depends entirely upon the Pc level, the ambient backpressure, and the expansion ratio; the only thing in it related to chemistry is the specific heat ratio, and that is near 1.2 for all the rocket combustion gases I ever heard of. CF is very sensitive to Pc level and expansion ratio. It also has the nozzle kinetic energy efficiency embedded in it, in the momentum term.

The strongest influences on Isp are c* (chemistry), CF (Pc and expansion), and BF (every % of bleed is a % right off the top).

Using the numbers just quoted, a BF = 0, and a testable-at-sea-level vacuum (!!!) expansion CF=1.80 at 2000 psia, we have LOX-LCO Isp = 4155*1.8 (1-0)/.995*32.174 = 4155*.0562269 = 233 sec, for LOX-RP1 Isp = 5900*.0562269 = 332 sec, for LOX-LCH4 Isp = 6120*.0562269 = 344 sec, and for LOX-LH2 Isp = 447 sec. (BTW, the CO value of 233 sec is less than the typical sea level (!!!) value of 250-255 sec for a high-solids, aluminized, AP-oxidized, composite solid made with HTPB or CTPB as the binder).

You only get the higher values folks often quote, if you push the state of the art in terms of nozzle expansion and chamber pressure. Mostly by higher chamber pressure, which is limited by the capabilities of the turbopump design. Which explains the push toward 4400 psia (300 bar) in things like SpaceX's Raptor, and why Raptor 1 never achieved intended performance.

One of the lessons in the course teaches how to use c* data and the CF result to get Isp, where the CF result comes from decisions about Pc, back pressure, separation, and expansion ratio. All of that plus a database for c* and r models vs propellant combinations, is in a spreadsheet supplied with the course.

GW

Last edited by GW Johnson (2023-11-13 09:26:29)

tahanson43206 · 2023-11-14 12:17:32

This is a demonstration of the enhanced powers of ChatGPT4 as of November, 2023...

Chat history
ChatGPT
Today
Patent Search Assistance
User
tahanson43206
ChatGPT 4
User
You
Patent Search Question for ChatGPT4 .... Hello Again ... this is a new topic ... you have new and very interesting powers, thanks to the hard work and generosity of your developers. Can you perform patent searches? I have a simple one in mind.
ChatGPT
Yes, I can assist you with patent searches. Please provide me with the details of the patent you're looking for, such as the patent number, inventor's name, company name, or specific technology, and I'll do my best to find the information you need.
User
You
I've already found this patent, so I know it exists, but I misplaced the reference. the inventor is GW Johnson of McGregor Texas. The subject matter is "cactus" and the application is removal or remover. There are a surprising number of patent holders with the last name Johnson and the initials GW. This one has the first name of Gary, but there are several of those. You can eliminate all the candidates who hold patents on subjects other than "cactus". You should find only one patent by that criteria.

ChatGPT
From a quick search, here's what I found.
I found the patent you were referring to. It is a patent granted to Gary W. Johnson for an "Apparatus for mechanically controlling and eradicating cactus and other succulent plants from farm and ranch pastures and other lands." This patent, numbered 6722446, was filed on March 25, 2002, and granted on April 20, 2004. The invention describes a machine designed to control and eradicate cactus and other succulent plants by using a cutter deck to break off cactus pads near ground level and a crushing rail to crush and mangle cactus pads and large portions of other succulents, thus preventing them from re-rooting.
User
You
Bravo! Nice work! I'm impressed and appreciative of the hard work your designers put into giving you this capability. Very Nice!
ChatGPT
Thank you very much for your kind words! I'm here to help with any other questions or searches you may have. Don't hesitate to ask!
Message ChatGPT…

There may be other patent holders in the NewMars membership of over 1000 folks. I'd be happy to show their patents if they let us know they exist.

(th)

GW Johnson · 2023-11-15 10:25:42

Hey, you found it!

I have another one, much older. US 4,416,112 from 1981. Titled "Fuel Injector for Ducted Rocket Ramjet Motor". I modified and adapted a 5-ported showerhead injector invented at UTC-CSD for a fixed-flow gas generator-fed ramjet to work with the throttle valve on our variable-flow gas-generator-fed ramjet. Yes, we really did fit a solid propellant device with a variable throat area throttle. No, CSD's fuel injector would not work on our throttle valve, but my rather serious internal modifications to it did, brilliantly.

Just to let you know, I made the best guesses I could for c* and r vs Pc models for LOX-LCO from the documents Kbd512 provided, and I added it to the database in the rocket performance spreadsheet for the course. I also found and corrected an error in the r-model in the spreadsheet. It does not affect Isp, but it does affect the oxidizer-fuel split for estimating tank volumes. I guess I need to send you the revised copy to replace the older one.

GW

tahanson43206 · 2023-11-29 11:18:20

For GW Johnson....

RobertDyck wrote:

Ps. Re post #42: Wonderful details! I bow at the feet of the master.

I don't recall seeing praise at that level in any of the thousands of posts I have seen in this forum.

It is unlikely to occur again, for a very long time.

This outfit is staffed with tough graders.

Congratulations for this ephemeral moment.

(th)

tahanson43206 · 2023-12-09 12:15:34

GW Johnson has completed preliminary studies of a variety of options for Lunar flights.

The best one is a bit of a surprise .... a vehicle using LH2/LOX does best in a direct flight from the Lunar South Pole (were there is thought to be plenty of ice) to LEO.

This space is reserved for a link to the pdf ...
https://www.dropbox.com/scl/fi/swegj9gd … brrpt&dl=0

(th)

GW Johnson · 2023-12-10 09:16:27

The link works, and I can see the paper that I wrote just fine. Thanks, Tahanson43206.

GW

GW Johnson · 2023-12-10 10:38:13

I found a small news item on the NBC News site that says the X-37B is going to be launched with a Falcon-Heavy rocket. That launch has been delayed until tomorrow (Monday). It's always been an Atlas-5 before.

BTW, the heat shield on X-37B is about the same as that on the space shuttle, except: they are using a two-piece ceramic refractory tile denoted "Tufroc" at the stagnation zones, instead of carbon-carbon ablative.

GW

tahanson43206 · 2023-12-11 11:55:16

This post is for two files on Lunar Space Vehicles and Flight Plans ...

These files contain corrections to the same files published earlier...

PDF:
https://www.dropbox.com/scl/fi/kzej67iq … t0bq7&dl=0

PowerPoint:
https://www.dropbox.com/scl/fi/ltrgyxhj … yk1nv&dl=0

(th)

tahanson43206 · 2023-12-12 11:20:44

For GW Johnson ...

Our contact in Houston (at NSS North Houston) just called ...

If I understand correctly, he has forwarded all four of your videos, but (as we have discussed) the recipient is buried in responsibilities.

Our contact allowed me to (at least attempt to) persuade him to set up a YouTube channel. I did that years ago, and things have changed a bit, but most of the procedure is the same or similar.

He ** should ** be able to post your videos.

As a bonus, our contact is a member of a Facebook group with (about) 300 real people members which has a shared interest in space news. He said he would think about posting announcements there about your videos, when he has them posted.

(th)

tahanson43206 · 2023-12-16 08:50:32

For GW Johnson ...

This post is about the sloshing problem that is in discussion in other topics....

There is a type of fishing mechanism that involves a one-way travel path.

I looked a bit further and found a system called a "Check Valve" that is used in industry.

Per Google:

Check valves are one-way or nonreturn valves that open with fluid movement and pressure, and close to prevent backflow of the pressure to upstream equipment such as pumps and compressors. In fact, check valves allow the fluid to flow in only one direction.
Check Valve - an overview | ScienceDirect Topics
www.sciencedirect.com › topics › engineering › check-valve

It seems to me your objection to SpaceNut's idea of separate tanks for the stage separation was that any tank that is holding liquid needs expansion room, so such a tank will always be subject to sloshing. However, a "Check Valve" design would not have that problem.

A "Check Valve" tank design would allow liquid to flow down the tank toward the engines during normal flight, but would close and prevent movement of the liquid back up the tank in the event of negative G force on the liquid.

Such a "Check Valve" could be made of metal, so there would be no issues of bladder failure due to low temperatures.

(th)

New Mars Forums

Announcement

#176 2023-10-01 12:10:20

Re: GW Johnson Postings and @Exrocketman1 YouTube videos

#177 2023-10-01 12:25:57

Re: GW Johnson Postings and @Exrocketman1 YouTube videos

#178 2023-10-01 20:09:59

Re: GW Johnson Postings and @Exrocketman1 YouTube videos

#179 2023-10-04 07:25:25

Re: GW Johnson Postings and @Exrocketman1 YouTube videos

#180 2023-10-04 08:57:55

Re: GW Johnson Postings and @Exrocketman1 YouTube videos

#181 2023-10-27 09:20:57

Re: GW Johnson Postings and @Exrocketman1 YouTube videos

#182 2023-10-29 10:08:43

Re: GW Johnson Postings and @Exrocketman1 YouTube videos

#183 2023-10-29 20:35:37

Re: GW Johnson Postings and @Exrocketman1 YouTube videos

#184 2023-10-30 18:23:52

Re: GW Johnson Postings and @Exrocketman1 YouTube videos

#185 2023-10-31 08:43:38

Re: GW Johnson Postings and @Exrocketman1 YouTube videos

#186 2023-11-07 11:10:58

Re: GW Johnson Postings and @Exrocketman1 YouTube videos

#187 2023-11-10 15:12:03

Re: GW Johnson Postings and @Exrocketman1 YouTube videos

#188 2023-11-12 20:19:13

Re: GW Johnson Postings and @Exrocketman1 YouTube videos

#189 2023-11-12 22:52:50

Re: GW Johnson Postings and @Exrocketman1 YouTube videos

#190 2023-11-12 23:36:18

Re: GW Johnson Postings and @Exrocketman1 YouTube videos

#191 2023-11-13 09:15:01

Re: GW Johnson Postings and @Exrocketman1 YouTube videos

#192 2023-11-14 12:17:32

Re: GW Johnson Postings and @Exrocketman1 YouTube videos

#193 2023-11-15 10:25:42

Re: GW Johnson Postings and @Exrocketman1 YouTube videos

#194 2023-11-29 11:18:20

Re: GW Johnson Postings and @Exrocketman1 YouTube videos

#195 2023-12-09 12:15:34

Re: GW Johnson Postings and @Exrocketman1 YouTube videos

#196 2023-12-10 09:16:27

Re: GW Johnson Postings and @Exrocketman1 YouTube videos

#197 2023-12-10 10:38:13

Re: GW Johnson Postings and @Exrocketman1 YouTube videos

#198 2023-12-11 11:55:16

Re: GW Johnson Postings and @Exrocketman1 YouTube videos

#199 2023-12-12 11:20:44

Re: GW Johnson Postings and @Exrocketman1 YouTube videos

#200 2023-12-16 08:50:32

Re: GW Johnson Postings and @Exrocketman1 YouTube videos

Board footer