Talk:Single-stage-to-orbit

Space Gun/Supergun/Gerald Bull

The idea of a space gun deserves some real estate if you ask me.

anonymous 15:10, 5 June 2010 (UTC) —Preceding unsigned comment added by 206.125.93.234 (talk)

I added a remark.--Patrick (talk) 06:46, 6 June 2010 (UTC)

No, it doesn't. A projectile fired from a "space gun" is not a rocket, and has no business in a discussion of SSTOs. Voronwae (talk) 00:31, 27 June 2012 (UTC)

DC-X

The first part of this is about the DC-X experimental vehicle. This is an interesting topic and Wikipedia should have an article on it. But why is it HERE? The connection of DC-X to single stage to orbit is rather tenuous. DC-X never made orbit and was not intended to. The intention was to test new technologies that are rather neat, but have no obvious connection to SSTO.

You could say the same about the X-33. The vehicle never made orbit and was not intended to. The X-33 was to have been a subscale prototype for a later vehicle.WolfKeeper 15:10, 29 August 2007 (UTC)

DC-X was a subscale prototype for a follow-on vehicle called DC-Y. DC-X was intended to pipeclean technologies such as VTVL, and fast turnaround with small crews, that the DC-Y would use. Whether DC-Y would have worked is a different question (whether any SSTO would work is a different question).WolfKeeper 15:10, 29 August 2007 (UTC)

The article doesn't even define what single stage to orbit means. It only briefly and peripherally explains why it might be a good idea. It doesn't discuss any of the difficulties with the concept, and there certainly are difficulties, at least in the minds of the guys who sign the checks. We have being doing orbiters for 40 years with a great variety of hardware, and NONE of it was SSTO.

For all these very serious criticisms i get the feeling the author still knows more about the subject than i do, so i'm not going to jump in and fix it just yet. But someone, please do.

Definition

Material from SSTO, take what you need:
Short for Single Stage to Orbit. In contrast to multistage rockets, the whole vehicle reaches orbital velocity. It is believed that SSTOs would led to much reduced costs for the access to space and allow aircraft like operations. The main problem in constructing such a vehicle is to make the engine efficient and the vehicle structure lighweight enough to avoid carrying excessive amounts of fuel or to be forced to drop away parts of your rocket while in flight. All attempts in constructing such a vehicle (DC-X, Roton, X-33) have so far been unsuccessfull due to technical and/or economic difficulties.

SpaceShip One

I do not see any discussion of Space Ship One...

At this point, SS1 is neither Single-Stage (carried initially by White Knight aircraft) nor Orbital. Lomn 21:38, 2 August 2005 (UTC)

Article needs rewriting

Appreciate the effort that's gone into this article, but it has quite a few grammar and writing style problems. Some Wikipedia style guidelines: [[1]]. There are also several misleading or speculative statements not appropriate for an encyclopedia. Examples:

"The lack of such abort modes on the Shuttle requires incredible failure avoidance costs and massive overhauls", given in the subheading "Why SSTO". This is wrong for several reasons: (1) The shuttle has many abort modes (2) The relationship of abort modes to SSTO is tenuous and shouldn't be listed in this subheading. The conventionally staged Soyuz is listed, which is a non sequitur, tangent to the subheading topic.

Another example: several "conventional wisdom" statements about the space shuttle are speculative and of uncertain veracity. E.g, "...high cost per launch of the Space Shuttle (a vehicle ironically designed to reduce launch costs)." Testimony of key shuttle officials during hearings by Columbia Accident Investigation Board indicates that's not the case. Operational costs were not unexpectedly high, but largely in line with early estimtates. For an actual transcription of relevant testimony, see [2]. There is a difference between conventional wisdom popularly repeated in the press and actual facts. Encyclopedias should strive to stick to the facts.

"The final vehicle required massive amounts of maintenance after every launch. This shift was partly a result of the removal of various abort systems, requiring the vehicle to be made safe via intensive inspection." This seems speculative. What abort systems were removed? What is the source for stating the lack of abort systems is directly related to operating costs? How would an abort system (e.g, launch escape) have sufficiently decreased inspection requirements to make a significant difference in operating costs?

"The engines are removed and rebuilt, large amounts of the structure are taken off for testing, and the entire refurbishing cycle takes months." The engines are not rebuilt after each launch, they are inspected.

I'd also suggest a link to Tsiolkovsky rocket equation, along with simplified explanatory verbiage, as this is the underlying technical reason why SSTO is difficult.

I'll be happy to rewrite the article if you agree and think it would be beneficial. My goal would be to retain most of the current material, better adherence to Wikipedia style guides, elimination of problem areas like the above, and addition of some clarifying material. Joema 17:19, 29 December 2005 (UTC)

This page still needs very badly to be scoured for erroneous discussions and information. I landed on it by accident and after just giving it a quick glance I can see several problems. My impression is that the creator of the page is a space enthusiast who doesn't necessarily know the subject area very well but would like to learn. I don't have the time right now to go over it, but a fair amount of material really needs to be discarded.

Just off the top of my head, I'll point out that this sentence:

"Several research spacecraft have been designed and partially or completely constructed, including Skylon, the DC-X, the X-33, and the Roton SSTO. However, despite showing some promise, none of them has come close to achieving orbit yet due to problems with finding the most efficient propulsion system.[1]"

is incorrect. In the case of Skylon, it's really only in the beginning stages of development. DC-X was not an SSTO. X-33 was not an SSTO either, and failed because the people casting the composite tanks were not experienced at composites and were asked to do something they didn't know how to do. The Roton failed because the economy crashed and the investors pulled their funding.

There are a lot of problems like this one throughout the article, in addition to some grammatical problems. Needs work! Voronwae (talk) 00:44, 27 June 2012 (UTC)

Article now rewritten

I've essentially rewritten the entire article to improve readability, remove numerous inaccuracies and POV statements. Encyclopedia articles should be non-judgmental and mainly descriptive information about the topic. Following the principles in The Elements of Style will greatly improve clarity. It's especially important to avoid unnecessary words and sentences. Don't use passive voice, and avoid long prepositional phrases. To quote The Elements of Style: "A sentence should contain no unnecessary words, a paragraph no unnecessary sentences, for the same reason that a drawing should have no unnecessary lines and a machine no unnecessary parts." See http://www.crockford.com/wrrrld/style3.html#13.

Russian speakers claim previous sentence contains half dozen unneeded words. --66.41.154.0 (talk) 01:32, 3 July 2013 (UTC)

More work needed

I agree with Wolfkeeper about article needing more work. Starting that work now.... Joema 22:51, 9 February 2006 (UTC)

As suggested, I've made numerous improvements. Reducing word count improves readability, yet doesn't require sacrificing meaningful content. See Elements of Style. In particular, see: [3].

Other contributors: please discuss here any issues with the changes. Joema 17:25, 10 February 2006 (UTC)

Dense vs Hydrogen Fuels

"The end result is the thrust/weight ratio of hydrogen fueled engines is 30-50% lower than comparable engines using denser fuels.

This inefficiency indirectly affects gravity losses as well; the vehicle has to hold itself up on rocket power until it reaches orbit. The lower thrust of the hydrogen engines means that the vehicle must ascend more steeply, and so less thrust acts horizontally. Less horizontal thrust results in taking longer to reach orbit, and gravity losses are increased by at least 300 meters per second. While not appearing large, the mass ratio to delta-v curve is very steep to reach orbit in a single stage, and this makes a 10% difference to the mass ratio on top of the tankage and pump savings." Shouldn't the hydrogen vehicle be the one that actually has a higher t/w ratio and lower gravity losses? Even if the engines have 30-50% lower thrust, hydrogen is so much lighter than RP1 that there should be more than 50% reduction in the total vehicle mass.

As it is now, this section does not seem very NPOV. It concentrates on the advantages of kerosene while ignoring many of the advantages of hydrogen. It also seems to imply that some dense fuels are better hydrogen for SSTOs, even though many studies concluded that hydrogen is the best option.--Todd Kloos 05:07, 25 March 2006 (UTC)

[Disclaimer: IANA Rocket Scientist]

I believe the quote is correct[4]. H2/O2 has a higher I_sp, but higher ΔV to orbit.

—wwoods 09:54, 25 March 2006 (UTC)

Minor Clarification suggested

... the lower lunar gravity and relatively thin atmosphere makes this much easier than from Earth.

This sentence should be rephrased for accuracy - the moon has practically no atmosphere. According to the Wikipedia entry, lunar atmospheric pressure is only 3 × 10^-13 kPa. Calling this near vacuum a "relatively thin atmosphere" sounds like a tortured euphemism.

Textor 06:10, 25 April 2006 (UTC)

specific impulse

The section on dense vs. hydrogen fuels uses the units of 'seconds' in reporting specific impulse. The measure is of impulse per unit mass so the number is about 9.8 times higher in SI. The numbers given are for pounds-force seconds per pound mass. In SI the metric is Newton-seconds per kg. The article on 'specific impulse' is similarly unclear on this point. 4.232.3.182 17:02, 15 August 2006 (UTC)

Trolling can be fun can't it? Anyway, pounds-force seconds per pound mass is heavily deprecated here, and *very* rarely used elsewhere. Newton-seconds per kg (or m/s) is acceptable, when accompanied by the Isp in seconds. But seconds is actually correct in both metric units and non metric units, the defining equations can be found in most textbooks.WolfKeeper 18:32, 15 August 2006 (UTC)

Competent rocket engineers know that the definition of specific impulse is the impulse delivered per unit mass of propellant expended; thus the name. Competent rocket engineers also know that the correct units for specific impulse are lbf-sec/lbm in U.S. units and N-sec/kg in SI. If they are old enough they may even know how the erroneous units of seconds happened. 4.232.0.191 13:47, 22 August 2006 (UTC)

Ratios

The ratios discussed in the "Why SSTO?" section are incorrect. A 2 to 1 ratio would be 2 parts of "X" and one part "Y", not one-to-one as stated in the article. The other ratios are wrong, too. A 5 to 2 ratio is 28.6%, not 40%, and a 4 to 1 ratio is 20% not 25%. I would correct it, but I don't know the actual fuel to structure ratios of the aircraft. Does the stunt plane really have a 4 to 1 ratio structure/fuel ratio (20%) or is it 25% fuel (3 to 1 structure/fuel ratio)? Martylunsford 15:17, 30 October 2006 (UTC)

Dunno about those in particular, but you're definitely wrong about your ratios for this article. Check out mass ratio; it points out that in rocketry, the ratio that is important is m0/mf (m0 is the initial takeoff weight, and mf is the final weight when the rocket ends its burn). As you can see a 4:1 ratio (m0/mf) is 25% dry weight. That ratio is also important in aircraft for calculating range. In both cases the ratio is not the ratio of fuel to structure; that's mathematically awkward for rocketry and aircraft range calculations and is never used (not without an explicit label anyway).WolfKeeper 18:36, 30 October 2006 (UTC)

I changed it anyway, it's more accurate and less confusing to have dimensionless number. That's the way Sutton does it.WolfKeeper 18:55, 30 October 2006 (UTC)

Hydrogen peroxide for fuel?

How come nobody mention about H2O2 as fuel? TheAsianGURU 21:50, 11 September 2007 (UTC)

Probably because the performance of propellant combinations including H2O2 is too poor to be used for SSTO.

It has actually been seriously suggested, see http://web.archive.org/web/20011119185055/http://www.im.lcs.mit.edu/bh/analog.html and Black Horse (rocket). Martijn Meijering (talk) 23:00, 3 April 2016 (UTC)

conclusions section?

I'd like to see a summary or conclusions section if possible. The nearest I can find is a sentence in the "Problems" section that says it is not "clear if a reusable SSTO carrying a useful payload can be built and operated for a feasible cost with current or near-term projected technologies."

Can we have a clear conclusion in the intro or in a conclusions section that says "the engineering challenges of developing a commercially useful SSTO launch system have to-date proven insurmountable" and perhaps an explanation to the effect that "SSTO's theoretical margins for feasibility are thin in any case"?

I understand that some people believe that SSTO could have gone somewhere if the government, in the form of NASA, wasn't in there mismanaging it. But I think that as a neutral encyclopedia we cannot assume that, and must rather take the failed efforts to date, like the X-33, at face value as sincere attempts that just couldn't solve the problem.Bdell555 (talk) 06:47, 26 July 2008 (UTC)

The thing is, Skylon's margins don't seem that thin at all- the projected payload fraction is about twice that of the Shuttle.- (User) WolfKeeper (Talk) 14:15, 26 July 2008 (UTC)

But you can say anything in the wikipedia you can reference.- (User) WolfKeeper (Talk) 14:15, 26 July 2008 (UTC)

Oxidizer weight

The article seems limited in the scope of ideas being investigated (even just for propulsion systems).

Re: "Oxidiser tanks are very lightweight when empty, approximately 1% of their contents, so the reduction in orbital weight by airbreathing is small, whereas airbreathing engines have a poor thrust/weight ratio which tends to increase the orbital mass."

This paragraph is misleading because it completely ignores the weight of oxidizer itself.

If for example, the weight of the oxidizer tankage was say 5% of the total rocket structure weight (in reality its probably more), then the weight of the oxidizer can be found using the following rationale (based on the statement in the article):

tankage=1/100*oxidizer

multiply both sides by 100: oxidizer=100*tankage

tankage=5/100*structure

substitute: oxidizer=100*5/100*structure=5*structure

So from this rationale, the weight of the oxidizer tankage may be 1% of its contents, but the weight of the oxidizer itself is 5 times the total structural weight of the rocket!

Why carry all this weight from a standstill when its all around us for the portion of the trip that requires the most thrust? Much of the fuel and oxidizer carried in conventional rockets is expended before reaching altitudes and speeds at which airbreathing engines become useless, because in this phase you're carrying the most fuel, gravity is strongest, drag is highest, and you have to start from stationary and get off the launchpad (inertia).

—Preceding unsigned comment added by 203.129.23.146 (talk) 08:09, 3 September 2010 (UTC)

Launch Assists

I here have a question: From my understanding, an SSTO needs to be single staged, assistance rather renders this not correct, especially as in the cases of air launches, but probably also sled launches, mass drivers, 'single stage to tethers' et cetera. It should be mentioned that these are not true SSTO's, or the section should be seriously altered or removed, — Preceding unsigned comment added by 97.117.238.96 (talk) 05:27, 12 February 2013 (UTC)

This Page is in dire need of a rethink!

A lot more work and research needs to be done in order to make this page up to the standard that it could be!

I edited out the mention of the Lynx spacecraft as an SSTO- its not, I believe it serves a similar role as the Virgin galactic spaceship. Please do the right research rather than making things up. — Preceding unsigned comment added by 82.33.168.225 (talk) 02:28, 31 July 2013 (UTC)

Also much out of date information — Preceding unsigned comment added by 82.33.168.225 (talk) 02:30, 31 July 2013 (UTC)

You are enthusiastically invited to help improve the page! Update the out-of-date or incorrect information, and say where you got the correct information using references. VQuakr (talk) 04:06, 31 July 2013 (UTC)

OTRAG Section

Hi

I have removed the following section on the OTRAG rocket; "and might have been successful if the project was not killed following political pressure from France, the Soviet Union and other parties." not only is it completely unsourced like the vast majority of this section but it also seems to smell a tad POV by using the term 'killed', surely 'cancelled' or 'shelved' is more apt. Feel free to revert if this was the wrong course of action however please add a fact tag if you do as the OTRAG article barely mentions meddling external forces being responsible and even when it does they are also completely unsourced too. Mishka Shaw (talk) 19:24, 3 April 2014 (UTC)

Space Shuttle Marginality Section

The third paragraph mentions that the Space Shuttle provided 96% of the needed energy to enter orbit from just the ET and main engines. The wording implies that the ET and engines alone could carry the Shuttle 96% of the way to orbit(i.e. a launch without the SRBs). But the Shuttle wouldn't have enough thrust to even take off. The SRBs provided a boost in more than just Delta-V, meaning that they get the Shuttle off of the ground and give it enough velocity to complete the main burn before reaching apoapsis, but the paragraph doesn't mention this. — Preceding unsigned comment added by 69.130.245.206 (talk) 18:10, 18 June 2014 (UTC)

"The marginality of SSTO can be seen in the launch of the space shuttle" - I disagree with the supporting arguments. I see a fatcs flaw so big this statement should be removed.

The current article text states that 4% of the energy needed to reach orbit was contributed by the first stage. This gives the impression that the second stage did the lion's share of the work and that if earth gravity was a little weaker or if rocket fuel was a little stronger that then a multistage rocket would not be needed.

I claim that the shuttle first stage does the lion's share of the work in direct opposition to the article. Further that for the specific example of the shuttle, a correct examination of the facts reveals that achieving orbit in a single stage is well out of the realm of the possible and not at all marginal as stated.

The shuttle's first stage expends 82% of the total vehicle fuel. (SRBs have 1,100,000 pounds of fuel each and it is all consumed during the first stage burn. The liquid tank uses about 20% of its 670,000 pounds of fuel during first stage burn. ) The current article suggests that only 4% of the energy needed to get to orbit was provided by the first stage. Well, by looking at the fuel staging attribution, you can see that the first stage provided 82% of the energy and not 4%.

The author's flaw was that he or she only examines gravitational potential energy and kinetic energy. ( g.p.e & k.e. ) There is an enormous amount of additional energy needed to reach orbit ( which never becomes g.p.e. nor k.e. ) This additional and extremely significant energy requirement was omitted from the article. I will call the omitted energy, 'parasitic station keeping' energy. In the role of 'parasitic station keeping' a very significant portion of the thrust is used solely to directly oppose the pull of gravity less the pull of centrifugal force. As the powered ascent progresses the C.F. increases over time of the burn, to match the gravitational force at which point orbit is reached and no more 'parasitic station keeping' energy is squandered.

During ascent the velocity vector and the thrust vector significantly diverge so that a component of the thrust vector can oppose gravity less C.F. That is the unfortunate expenditure of much energy which is never realized as g.p.e. nor k.e. To illustrate further, imagine a rocket producing one g of force. It would remain motionless. It would burn through all its full and none of the energy from its fuel would become g.p.e. nor k.e. Now imagine a rocket producing 1.1G of thrust. Again the rocket would burn through its fuel and only a tiny tiny fraction of its fuel would become g.p.e. or k.e. the rocket would never reach orbit. This is the inefficient consequence of having a slowly accelerating rocket. ( a manned rocket for example ) A rocket must burn its fuel in as short a period of time as is feasible to achieve the best efficiency.

For unmanned and for more structurally robust rockets a higher G force is employed. This has the effect of greatly reducing the parasitic station keeping requirement and transforming the rocket's energy more efficiently into g.p.e and k.e. Only if a rocket could have an instantaneous burn of all its fuel could it avoid the parasitic station keeping. In such a fictional circumstance the article would be more roughly sound in calculating the needed energies to reach orbit as the sum of only kinetic and g.p.e.

The article claim that single stage to orbit is marginally obtainable may deserve an exploration but the example of the space shuttle may well have been the worse possible choice to support this claim. The space shuttle example does the opposite. It destroys the claim. If the marginality of s.s.t.o. claim is to stay in the article it should be pointed out that it is only true for rocket configurations that have the highest of acceleration profiles. ( rugged and unmanned ) Further that as the acceleration profile weakens ( eventually to 3Gs for manned flights such as with the shuttle ) the ability to achieve orbit in a single stage is well out of the realm of the possible and just the opposite of marginal. — Preceding unsigned comment added by 99.233.79.170 (talk) 07:00, 12 February 2015 (UTC)

True, and this isn't even taking into account that the first stage has to go through the most dense part of the atmosphere, where you necessarily waste energy, by some combination of "parasitic station keeping" and atmospheric drag (thus a high acceleration does not solve you problems). It is unfair to compare the delta-v of the stages without taking this into account. For now, I removed the sentence in question, though I am open to adding a comparison of SSTO to the Space Shuttle, if it can be supported by citations. --Tobias (Talk) 04:41, 8 April 2016 (UTC)

tri-propellant

is it worth adding a section on tripropellant fuel? — Preceding unsigned comment added by Patbahn (talk • contribs) 07:14, 21 November 2015 (UTC)

Alternative approaches: big dumb boosters

"This is somewhat similar to the approach some previous systems have taken, using simple engine systems with "low-tech" fuels, as the Russian and Chinese space programs still do."

Howso? Why should kerosene/lox & N2O2/UDMH be considered low-tech propellants? Moreover, the Russians introduced the oxidizer-rich staged combustion cycle in the 60s and used it in the NK15, NK33 , RD253, RD120, RD170/171 and the derived RD180, RD151, RD181, RD191, et al, several of which remain in use, some in western launch vehicles. These are complex engines with performance yet to be attained by western engines with the same propellants. Western high performance engine designs relied on LH2/LOX propellants instead, which the Russians eschewed for the same reasons mentioned earlier in the article. In the Chinese Long March launch vehicle family, the new YF-100/115 series lox/kerosene engines are also oxidizer-rich staged combustion designs. LH2/LOX upper stage engines (YF-73/75/77) have been used in the Long March 3 and its descendants. — Preceding unsigned comment added by Jmostly (talk • contribs) 10:19, 4 April 2016 (UTC)

"Low-tech" apparently is someone's glib way of saying low-specific impulse. JustinTime55 (talk) 16:05, 4 April 2016 (UTC)

should this point to it's own article to keep the main article on point? --Patbahn (talk) 02:58, 21 February 2019 (UTC)

SpaceX's BFR / BFS

It might be worth mentioning the SpaceX's BFR on the page, since Elon Musk has stated a few times recently that the BFR's second stage (BFS) is capable of getting into low earth orbit by itself without needing the 1st stage booster, most recently at the Falcon Heavy post-launch press event: https://www.youtube.com/watch?v=F7mw2_pfcz4 ("The ship is capable of single stage to orbit if you fully load the tanks." 50:00)

Although he also said a few months earlier in a Reddit AMA that: "Worth noting that BFS is capable of reaching orbit by itself with low payload, but having the BF Booster increases payload by more than an order of magnitude. Earth is the wrong planet for single stage to orbit." - which sounds like although it might be possible, it might not be something they see as viable depending on how they define "low payload". Unless plans have changed since then of course.

82.15.131.175 (talk) 02:53, 8 February 2018 (UTC)

It might be worth explaining how this is possible (probably with insufficient fuel left for a reentry and landing) - but there seems no intention to actually do it. - Rod57 (talk) 10:11, 15 July 2018 (UTC)

BFS planned to have a fuelled mass (without payload) of 1,185 tonnes. With 4 vacuum Raptor engines and 3 sea level Raptors ("The sea-level model Raptor engine ... is expected to have 1,700 kilonewtons (380,000 lbf) thrust at sea-level with an Isp of 330 s increasing to an Isp of 356 s in the vacuum of space. The vacuum model Raptor,... is expected to exert 1,900 kN (430,000 lbf) force with an Isp of 375 s.) We'd have to estimate the thrust and ISP of the vacuum engine at sea level. BFR (rocket) says BFS has 12.7 MN (2,900,000 lbf) total thrust, empty mass 85,000 kg (187,000 lb), fuel load 1,100 tonnes (2,430,000 lb). (Thrust/initial-weight ~1.1) structural coefficient ~ 0.072 so looking at the graph shows how critical the varying ISP is to attempt SSTO - Rod57 (talk) 18:38, 15 July 2018 (UTC)

Proposed talk page edit to delete External links modified sections

Per ongoing discussion at Wikipedia:Village_pump_(proposals)#RfC:_Delete_IABot_talk_page_posts? and Template_talk:Sourcecheck#Can_we_change_the_standard_message_to_says_its_OK_to_delete_the_entire_talk_page_section I'd like to delete the above External links modified section(s). Any objections ? - Rod57 (talk) 10:15, 15 July 2018 (UTC)

 Done - Rod57 (talk) 13:42, 19 December 2020 (UTC)

altitude compensation

perhaps the section on altitude compensation should be improved? --Patbahn (talk) 02:57, 21 February 2019 (UTC)

Starship SSTO

So SpaceX actually considers doing SSTO with Starship. https://www.nasaspaceflight.com/2019/05/spacex-ssto-starship-launches-pad-39a/ --91.79.173.55 (talk) 21:52, 17 May 2019 (UTC)

design challenges inherent in SSTO

The section closes with the conclusion "Given that current materials technology places a lower limit of approximately 0.1 on the smallest structural coefficients attainable,[30] reusable SSTO vehicles are typically an impractical choice even when using the highest performance propellants available."

the section is poorly cited, which is problematic but more problematic is that it's not factually well grounded. The Atlas 1, Centaur and Titan II core stage all had structural coefficients significantly below 5%.

--Patbahn (talk) 04:10, 31 May 2019 (UTC)

This section comes from Pg 631 of Curtis's book but it's not well grounded.

--Patbahn (talk) 14:27, 31 May 2019 (UTC)

Sounds like it is cited then? None of the three stages you mention are reusable (or even complete vehicles). VQuakr (talk) 15:35, 31 May 2019 (UTC)

The fact a stage can approach 95% structural margin argues the highest potential margin is 95%. Now, there is mass budget for reusable systems, but, your argument seems based on original research when the conclusions don't support that. --Patbahn (talk) 05:33, 26 January 2020 (UTC)

It is worth noting the Atlas 1 was capable of making orbit on a single stage, if you look at the Mercury-Atlas launch John Glenn flew that was 1 Stage to orbit, discarding the

booster engines in flight, rather then staging the whole stage off.--Patbahn (talk) 17:39, 22 March 2021 (UTC)

Same engine for all altitudes is just aerospike

There are no serious ideas for engines that maintain efficiency at all altitudes other than aerospikes that I know of. There are very few citations to justify this claim and the other articles about non-aerospike designs that this article links to have a similar lack of sources. Let's just remove the altitude adapting stuff and focus solely on aerospikes. Titan(moon)003 (talk) 19:00, 4 October 2024 (UTC)