Blue Origin photo

[no_way]

1 ) Redundancy and error recovery with chutes usually means cutaway and a reserve deployment. With control system problems, it usually means switching to a backup if you have one. <br />Backup control weighs (and costs ) less than a reserve.<br /><br />2) BO apparently plans chute only for a backup landing. You are arguing that they should be used for primary landing. Apples and oranges on an operational reuseable vehicle.<br />In my book, having a backup system for extra safety is completely OK and adding it is a tradeoff between payload and safety factors. The weight hit that you take is for extra redundancy.<br />Your argument that "having a chutes on board" and making use of them for landing .. light aircraft most of the time land on their wheels, not on chutes. Although they have "both on board", wheels for takeoff and chutes for emergencies.<br /><br />Yes i am still arguing weight with you, just dont confuse backup systems ( weight is traded off for extra safety and redundancy ) with primary methods of operations. See, if BO had chosen chutes as a primary landing method, theyd still have to have the extra one on board to achieve the same level of redundancy, which means that the overall weight penalty would be higher.<br />If you wanna get the flight prices down, you optimise your primary operations for as fast turnaround, simplicity and low maintenance as possible. Fuel tanks you fill up anyway between flights, pumping some extra for powered landing basically adds no operational complexity. <br />Packing or installing a new chute does add lots of operational complexity.<br />There is one more factor that makes chutes lose out, weather. Powered landing is obviously more tolerant of bad weather conditions.<br /><br />Regarding reserve chutes, i personally would add one extra layer of redundancy. Wearing a personal parachute and bailing out with that if the need should arise. Vostok style.<br /><br />BTW, there is no evangelical powered landing cult. I am just pointing out t

 

[gunsandrockets]

"Some facts for you:" <br /><br />"...Spacecraft on re-entry from orbital velocities enter the atmosphere almost at a tangent, travelling a considerable distance around the globe as they burn off velocity in the form of heat. A re-entry trajectory that travels around a third of the circumference of the Earth will travel through over 8,000 miles of atmosphere..."<br /><br />8,000 miles?<br /><br />A few facts about atmospheric reentry<br /><br />The planned reentry path of the Genesis sample return capsule was only 600 miles long...<br /><br />http://science.nasa.gov/headlines/y2004/03sep_genesisreentry.htm<br /><br />"...At precisely 8:52:46 a.m. Pacific Daylight Time (PDT), northwest of Bend, Oregon, a fireball will appear: a white-hot dot of light, brighter than the planet Venus, gliding across the blue morning sky..."<br /><br /><br />"...From Bend, traveling 25,000 mph (11 km/s), the fireball will streak across eastern Oregon, brightening as it descends into denser parts of Earth's atmosphere. At 8:53:35 a.m. PDT it crosses the southwestern corner of Idaho and, moments later, northern Nevada not far from the tiny town of Elko. Finally, at 8:54 a.m. PDT, slowed to a near-halt by the capsule's drogue parachute, the fireball will fade over Utah."<br /><br /><br />Even a spaceplane only takes half as much distance to reenter as the 8,000 miles you talked about. The Space Shuttle takes 3,700 miles from the point of atmosphere entry interface to the point of runway landing.<br /><br />http://www.spaceflight.nasa.gov/shuttle/reference/shutref/events/entry/<br /><br /><br />A few facts about recovery speeds<br /><br />The atmospheric terminal descent speed of a Blue Origin style VTVL RLV is determined by the drag of the body of the vehicle rather than the drag provided by a parachute. Since a RLV during

 

[gunsandrockets]

"Its about time." <br /><br />Here is a fascinating article that hobbyspace.com linked to...<br /><br />http://www.spacefuture.com/archive/history_of_the_phoenix_vtol_ssto_and_recent_developments_in_single_stage_launch_systems.shtml<br /><br />...it's not completely relevant to Blue Origin since Blue Origin will have to use a multi-stage rocket for orbital operations, but still an informative history of the evolution of VTVL RLV.<br />

 

[no_way]

<blockquote><font class="small">In reply to:</font><hr /><p>Since a RLV during landing is mostly a big lightweight almost empty fuel tank, there is a lot of drag a RLV can generate compared to it's weight. Some designs include drag flaps to increase vehicle drag even more. So a subsonic terminal descent speed is well within VTVL design possibility. <p><hr /></p></p></blockquote><br />Exactly. 10% fuel loaded New Shephard capsule's terminal velocity in lower atmosphere will be rather slow.<br />Try some calculations : <br />http://www.grc.nasa.gov/WWW/K-12/airplane/termv.html<br />

 

[mrmorris]

<font color="yellow">"Even a spaceplane only takes half as much distance to reenter as the 8,000 miles you talked about. "</font><br /><br />I didn't claim that Genesis (or any spacecraft) used that particular distance. I did a quick Google for the re-entry path and didn't find it, so I picked a fraction of the Earth's circumference off the top of my head and provided a distance to compare with the vertical thickness of the Earth's atmosphere.<br /><br />I do appreciate the two figures you provided. The Genesis return capsule at 225kg was able to largely shed its speed in only fifty times the vertical distance of the Earth's atmosphere, wheras the orbiters at ~80,000 kg require about 300 times the distance.<br /><br /><font color="yellow">"Since a RLV during landing is mostly a big lightweight almost empty fuel tank, there is a lot of drag a RLV can generate compared to it's weight."</font><br /><br />Keep in mind that we're talking about a suborbital craft rather than an SSTO. While an SSTO would indeed have 95+% of its mass in propellant, making it effectively a 'big balloon' on descent, the mass fraction on a suborbital will be *much* lower. Also -- the density of the H2O2/Kerosene mix that's an advantage on the way up is a disadvantage here -- smaller tanks, smaller craft, less drag. Earlier I calculated about 2,400kg of propellant/oxidizer for descent based on the thrust figures provided by BO. Assuming a vehicle mass of 2,400 kg -- that much propellant would provide a dv of 1,390 m/s (where my mach four figure came from). I thought (and think) a vehicle massing about 60% of Dragon seems somewhat reasonable, but we can assume a higher one, as I believe I indicated... ah, yes I did. If we assume NS will have a dry mass of 3400kg, the dv provided is 1,150 m/s (mach 3.3). If we assume NS will have a dry mass of 5400kg, the dv provided is 855 m/s (mach 2.5). <br /><br />If you truly believe that NS is going to be coming in at subsonic speeds

 

[mrmorris]

<font color="yellow">"You are free to start your own project and i'd cheer for you if you'd win next years lunar lander challenge with parachute landing craft. "</font><br /><br />I might have to try that out. If I could design a workable lunar lander that used a parachute, I'd soon gather a cult following of my own. That'd pretty much trump walking on water.<br /><br />Anyway -- I'm done with this thread.

 

[spacefire]

Glad you mentioned the parachute-rocket combo used by the Russians, I was thinking about it yesterday.<br />It makes most sense to have a 'soft' landing with a final rocket braking and a 'hard but survivable' landing by parachute alone.<br />But after the parachutes have opened, I don't think the passenger module can separate from the propulsion module. If the propulsion module malfunctions and the craft just plops down, it needs to protect its occupants.<br />That is, no fires due to the remaining propellant in the tanks.<br /> <div class="Discussion_UserSignature"> <p>http://asteroid-invasion.blogspot.com</p><p>http://www.solvengineer.com/asteroid-invasion.html </p><p> </p> </div>

 

[soyuztma]

<blockquote><font class="small">In reply to:</font><hr /><p>Keep in mind that we're talking about a suborbital craft rather than an SSTO. While an SSTO would indeed have 95+% of its mass in propellant, making it effectively a 'big balloon' on descent, the mass fraction on a suborbital will be *much* lower. Also -- the density of the H2O2/Kerosene mix that's an advantage on the way up is a disadvantage here -- smaller tanks, smaller craft, less drag. Earlier I calculated about 2,400kg of propellant/oxidizer for descent based on the thrust figures provided by BO. Assuming a vehicle mass of 2,400 kg -- that much propellant would provide a dv of 1,390 m/s (where my mach four figure came from). I thought (and think) a vehicle massing about 60% of Dragon seems somewhat reasonable, but we can assume a higher one, as I believe I indicated... ah, yes I did. If we assume NS will have a dry mass of 3400kg, the dv provided is 1,150 m/s (mach 3.3). If we assume NS will have a dry mass of 5400kg, the dv provided is 855 m/s (mach 2.5). <br /><p><hr /></p></p></blockquote><br />According to the EA New Shepard has 54,431 kilograms-mass of peroxide and kerosenen on board. If we take your dry mass number of 5400kg we have a total vehicle weight of 59831 kg. If we put these numbers in the rocket equation for a ve of 276*9.81=2708 m/s, we get a delta v of 6513m/s. This is way too much because when i calculated the required delta v for a suborbital vehicle i got something in the neighbourhood of 1000m/s, so i'm pretty sure the New Shepard will probably weigh over 20 metric tons dry. The New Shepard is a complete spacecraft and it is big: "The stacked vehicle would have a roughly conical shape with a base diameter of approximately 7 meters (22 feet) and a height of approximately 15 meters (50 feet)." You can not compare it to the Dragon.<br />You should really read the EA. Al <div class="Discussion_UserSignature"> </div>

 

[mrmorris]

OK -- you drew me back in. More numbers to crunch -- one of the many things I have trouble resisting.<br /><br />Answering one question in reverse:<br /><br /><font color="yellow">"I also want to know how you arrive at the 6,672,000 kg-m/sec number."</font><br /><br />My post here contains snippets from the EA which give a thrust level and burn time for the descent firing. I did assume a constant thrust in my calculations. In any event, I have nothing on which to estimate how much it trails off as the note you posted indicates it does. If they fire the landing thrusters for 15 seconds at 444,822 Newtons, the total impulse for landing is 6,672,000 kg-m/sec. That would require about 2,464 kg of N2O2/Kerosene assuming an isp of 276s. If the thrusters throttle down, then the total thrust used for landing will be an indeterminate amount less than this.<br /><br />OK -- refining the figures based on the numbers you provided:<br /><br />54,431 kg of H2O2/Kerosene at an exhaust velocity of 2708 m/s gives us a total thrust of 147,399,148 kg-m/s. <br /><br />Given the calculation for the propellant used to generate the landing thrust, let's simply assume that 52,000 kg even is used for ascent. Using the RE to calculate the final mass assuming a dv of 1020 m/s (mach 3 -- which is what SS1 got to approximately):<br /><br />1020 m/s = 2708 m/s * ln (Vf / Vf * 52000) - 9.8 m/s2 * 120 s <br />1020 m/s = 2708 m/s * ln (Vf / Vf * 52000) - 1176 m/s <br />2196 m/s = 2708 m/s * ln (Vf / Vf * 52000)<br /> 0.81093 = ln (Vf / Vf * 52000)<br /> 2.25 = Mf / Mf - 52000 kg<br /> Mf = 2.25 Mf - 117000 kg<br /> 1.25 Mf = 117000 kg<br /> Mf = 93,600 kg.<br /><br />This means the dry mass New Shepard is expected to be ~17% more than a shuttle orbiter!?<br /><br />If true -- the remaining 6,672,000 kg-m/sec of thrust would onl

 

[scottb50]

One possibility might be using a combined propulsion system, rockets and turbojets. Weight could be kept down by using the turbojets for initial launch and landing and the rocket motors for ascent only. I still question whether you could do it as a SSTO. <div class="Discussion_UserSignature"> </div>

 

[oker59]

i remember years ago, there was a new rocket technology development using turbojets that upped the performance of rocket technology higher than a long time.

 

[scottb50]

If the turbojets do nothing more than lift their own weight, and their fuel, they would be effective. When you figure the fuel and oxidizer required for a rocket motor it only makes sense to use turbojets when you do not need LOX. <br /><br />Where I have a question if it is worth it to land vertically. It would take a lot less fuel, for the turbo-jets to allow a conventional landing than a vertical landing and variables, weather especially would be less of a constraint. The other side would be having to have a landable, read winged, vehicle. As I see it wings would weigh less than the fuel required for a vertical landing, even if you had a powered approach and landing, with wings. <div class="Discussion_UserSignature"> </div>

 

[Boris_Badenov]

I suspect what they are really designing is an orbital vehicle. If that is the case wings are a liability. <div class="Discussion_UserSignature"> <font color="#993300"><span class="body"><font size="2" color="#3366ff"><div align="center">. </div><div align="center">Never roll in the mud with a pig. You'll both get dirty & the pig likes it.</div></font></span></font> </div>

 

[phaze]

My guess would be this is a design for the Moon/Mars.<br /><br />Vehicles like this could shuttle people to and from the surfaces and orbital platforms.

 

[soyuztma]

<blockquote><font class="small">In reply to:</font><hr /><p>1020 m/s = 2708 m/s * ln (Vf / Vf * 52000) - 9.8 m/s2 * 120 s <br />1020 m/s = 2708 m/s * ln (Vf / Vf * 52000) - 1176 m/s <br />2196 m/s = 2708 m/s * ln (Vf / Vf * 52000) <br />0.81093 = ln (Vf / Vf * 52000) <br />2.25 = Mf / Mf - 52000 kg <br />Mf = 2.25 Mf - 117000 kg <br />1.25 Mf = 117000 kg <br />Mf = 93,600 kg. <br /><p><hr /></p></p></blockquote><br />My calculations are purely for ascent.<br />The ascent delta v is arounde 2282 m/s: gravity losses are 9.81*110 ,from coasting from 38 km to 100 km we get 9.81*sqrt(2*62000/9.81) for a total of 2182 m/s. And we will include 100 m/s for drag losses.<br />For Isp 276 and a ascent delta v of 2282 i get a mass ratio of exp(2282/2708)=2.32 So that would mean a mass with landing propellants of 52/(2.32-1)=39.32 metric tons. And a total weight of 91.32 tons. <br />I also just found this paper: Flight Mechanics of Manned Sub-Orbital Reusable Launch Vehicles with Recommendations for Launch and Recovery It includes the necessary calculations. It also gives you a nice argument: the author claims that for powered vertical landing 20 to 30% of vehicle landing weight must be propellant. <img src="/images/icons/smile.gif" /><br />Astronautix has also posted there guess: New Shepard. But their total weight seem kind of low and their delta v rather high.<br /><blockquote><font class="small">In reply to:</font><hr /><p>Without more data and formulas than I have available, I can't disprove it, but I'm prefectly able and willing to have grave personal doubts.<p><hr /></p></p></blockquote><br />This seems to be the first reaction of a lot of people when they first hear about vertical landing: "wouldn't that consume a lot of fuel?" But people seem to forget about drag.<br />Maybe you should write your one little program to verify that it will be subsonic. <a></a> <div class="Discussion_UserSignature"> </div>

 

[Boris_Badenov]

<font color="yellow"> My guess would be this is a design for the Moon/Mars. </font><br /><br /> You could be right, & if that is the case it would mean cooperation & a division of effort that we are not hearing of in the media or from our "angel patrons"statements<img src="/images/icons/wink.gif" /><br /> SpaceX, RPK, SpaceDev or LockMart provide the transport to LEO.<br /> Bigelow has the LEO station. <br /> Bigelow & one of the above cooperate on a Lunar shuttle.<br /> Blue Origin provides the transport from the shuttle to/from the Lunar surface. <br /> A division of responsibilities like this would mean no one company would shoulder all the risk. That is just smart business. <div class="Discussion_UserSignature"> <font color="#993300"><span class="body"><font size="2" color="#3366ff"><div align="center">. </div><div align="center">Never roll in the mud with a pig. You'll both get dirty & the pig likes it.</div></font></span></font> </div>

 

[rocketman5000]

every company sholders all of the risk. You are placing other companies and their share holders in your critical path. if you are Bigelow and nobody gets to lunar orbit what good would your space station be? If you were BO with your speculated lunar lander and nobody go to LEO it wouldn't really mater if you had the best, cheapest and available lunar lander

 

[phaze]

things is... we have NASA/others to fill in the gaps if one or more of these companies were to fall behind.<br /><br />

 

[spacester]

I'm on the same page as Boris here.<br /><br />The point about critical paths is valid, but I would point out that a long time ago Intel put Microsoft on a critical path (somewhat on purpose and by accident) and that didn't work out too badly for either company.<br /><br />It's a critical path, which means the journey requires that the challenge be met and overcome. A critical path is not an insurmountable obstacle to walk away from. <br /><br />If all the alternative paths are not only just as critically difficult but require making the trip all alone, joining others on a common path makes a whole lot of sense.<br /><br />BTW, I dugg this story and it briefly made it to the front page. Way fun, I think I'll make a comment there referencing this thread. Good stuff here, it's been fun watching this thread develop. <br /> <div class="Discussion_UserSignature"> </div>

 

[scottb50]

I don't see the need for the aerodynamic design, especially for the moon. It looks to me like it is designed for re-entry in an atmosphere. <div class="Discussion_UserSignature"> </div>

 

[mrmorris]

<font color="yellow">"Maybe you should write your one little program to verify that it will be subsonic. ... If you build this program post the results "</font><br /><br />Nope -- I really don't care enough about suborbital flight to go to that level of effort. But the AIAA paper you posted a link to was interesting. I'll graciously concede that the two PhD's who wrote it <b>might</b> know more about the relevant equations and modeling than I do. <img src="/images/icons/smile.gif" /><br /><br />The relevant portions of the paper can be summed up in the two graphs below. Basically from ~120,000 feet to ~60,000 feet (~36,600m to ~18,300m) the craft undergoes a *lot* of deceleration, hitting about 5.7 Gs at ~27,400 m. More than I'd ever have guessed... obviously. <img src="/images/icons/smile.gif" />

 

[dreada5]

http://www.astronautix.com/craft/newepard.htm

 

[rocketman5000]

thanks for the link, clarifies alot that was discussed above in this thread.

 

[Boris_Badenov]

Does the New Shepard look like it's too big to be a suborbital ship to anybody besides me? It's designed to carry 4, is 50' tall, 23' in diameter & weighs in at more than 150,000#. <br /> This thing is a monster compared to the Dragon or the GX-3. <div class="Discussion_UserSignature"> <font color="#993300"><span class="body"><font size="2" color="#3366ff"><div align="center">. </div><div align="center">Never roll in the mud with a pig. You'll both get dirty & the pig likes it.</div></font></span></font> </div>

 

[rocketman5000]

Its not really a monster when you consider compared to the Dragon and GX-3 that it lands vertically and seems to be built with much higher safety margins. Couple those two things with the fact a much lower isp fuel is being used and the craft can get heavy pretty quickly.