New Mars Forums

Mars Insight scheduled to land near the equator already will be a great geology mission but the recent discovery of large water deposits near the equator on Mars raises the possibility it could also be a great astrobiology mission:

Water ice found near Mars’s equator could entice colonists and life-seekers.
Find also poses climate puzzle.
SCIENCEMAG.ORG

http://www.sciencemag.org/news/2017/08/ … fe-seekers

The full research article is available free, full-text online:

Equatorial locations of water on Mars: Improved resolution maps based on Mars Odyssey Neutron Spectrometer data.
Icarus, Volume 299, 1 January 2018, Pages 148-160.
Jack T.Wilson, Vincent R.Eke, Richard J.Massey, Richard C.Elphic, William C.Feldman, Sylvestre Maurice, Luís F.A.Teodoro
http://www.sciencedirect.com/science/ar … 3516306029

Mars Insight is scheduled to land in Elysium Planitia at 4°N 136°E:

Mars Insight.
https://en.wikipedia.org/wiki/InSight#S … ic_payload

Here is a map from the Icarus paper showing the regions of high water(ice) content:

Notice in the second panel in the image there is a region of high water content indicated in blue near the Mars Insight landing site. However, for Mars landers you can't get the landing position exactly right and there is unfortunately a larger area above and to the right of very low water content indicated in red near the landing site. The question is could they stick the landing near the high water region?

One of the Mars Insight instruments might be ideal for detecting the near surface water since it will include a subsurface borer able to drill down to 5 meters:

But what instruments could be used to distinguish liquid water from ice water?

In view of the possibility of liquid water existing near surface, I would also like to see a true microscope put on the lander. Of all the landers sent to Mars none included a true optical microscope. The best optical resolution of the imagers sent to Mars were no better than that of a geologist's hand lens.

   Bob Clark

You mean the "grain" simply means the weight of propellant?

  Bob Clark

Discussion of low cost orbital rockets for cubesats and smallsats that are within the capabilities of most university undergraduate labs:

Orbital rockets are now easy, page 2: solid-rockets for cube-sats.
http://exoscientist.blogspot.com/2017/0 … age-2.html

Feasibility comments encouraged.

    Bob Clark

I'll need some type of attitude control for the upper stages operating in little atmosphere, as fins would be ineffective. How did the hydrogen peroxide thrusters work?

  Bob Clark

Article doesn't say how much these motors repurposed as orbital launchers could loft to orbit. Anyone know that?

  Bob Clark

I found this commercial solid motor with a similar sea level Isp as the Cesaroni solid rocket motors Pro150 I mentioned earlier:

Star 37.
Thiokol solid rocket engine. Total impulse 161,512 kgf-sec. Motor propellant mass fraction 0.899. First flight 1963. Solid propellant rocket stage. Burner II was a launch vehicle upper stage developed by Boeing for the Air Force Space Systems Division. It was the first solid-fuel upper stage with full control and guidance capability developed for general space applications.
AKA: Burner 2;TE-M-364-1. Status: First flight 1963. Number: 180 . Thrust: 43.50 kN (9,779 lbf). Gross mass: 621 kg (1,369 lb). Unfuelled mass: 63 kg (138 lb). Specific impulse: 260 s. Specific impulse sea level: 220 s. Burn time: 42 s. Height: 0.84 m (2.75 ft). Diameter: 0.66 m (2.16 ft).

Thrust (sl): 33.600 kN (7,554 lbf). Thrust (sl): 3,428 kgf.
http://www.astronautix.com/s/star37.html

So we might estimate the vacuum Isp of the Cesaroni to be in the Star 37's range of 260 s.

Using this as a vacuum Isp of a first stage and a 285 s vacuum Isp of second and third stage I get a three-stage solid rocket able to reach orbit, with not even particular high mass ratios of 5 to 1 for the stages:

With a payload of 7 kg:

And this with standard nozzles, no altitude compensation required.

The Cesaroni Pro150 retails for about $3,000 and in general the Cesaroni solids cost in the range of $100 per kg of the motor mass:

Cesaroni O8000 White Thunder Rocket Motor.   
$3,099.95
Product Information
Specification
Brandname:  Pro150 40960O8000-P                 Manufacturer:  Cesaroni Technology
Man. Designation:  40960O8000-P                    CAR Designation:  40960 O8000-P
Test Date:  4/10/2008                                     
Single-Use/Reload/Hybrid:  Reloadable              Motor Dimensions mm:  161.00 x 957.00 mm (6.34 x 37.68 in)
Loaded Weight:  32672.00 g (1143.52 oz)         Total Impulse:  40960.00 Ns (9216.00 lb/s)
Propellant Weight:  18610.00 g (651.35 oz)       Maximum Thrust:  8605.10 N (1936.15 lb)
Burnout Weight:  13478.00 g (471.73 oz)          Avg Thrust:  8034.50 N (1807.76 lb)
Delays Tested:  Plugged                                    ISP:  224.40 s
Samples per second:  1000                               Burntime:  5.12 s
https://www.csrocketry.com/rocket-motor … motor.html

So take the cost of the third stage, derived from the Cesaroni Pro1050, as $3,000, and the second stage 4 times larger as $12,000, and the first stage larger by an additional factor of 4 as $48,000. So $63,000 for a small sat launcher with a 7 kg payload to orbit.

  Bob Clark

As described on the Astronautix.com page, the Soviet LK did not do the full landing on its own. It required the Soviet Block D stage to do most of the propulsion towards the lunar surface, bringing the LK to within 5 km and 100 m/s of the landing. The rest of the landing was done by the LK, as well as doing the take-off.

This type of landing used what's called a "crasher stage" since the Block D would crash into the lunar surface. I'm not a fan of this type of landing, though it's probably just a visceral reaction with this stage crashing and exploding on the lunar surface, even though kilometers away from the lander.

In any case the mass of the LK stage was comparable to the Apollo's ascent stage only, and the full lander was not at a saving in mass to the Apollo full lander.

About the Block D stage:

Block D space tug.
http://www.russianspaceweb.com/n1_d.html

For a replacement for this Block D stage, I like the storable propellant version of the Ariane 5 upper stage:

Ariane 5-2
N2O4/MMH propellant rocket stage. Storable propellant, restartable upper stage for use with Ariane 5. Chamber pressure 10 bar; expansion ratio 83.0; propellant mix ratio 2.05. Empty mass without VEB payload fairing support ring and avionics is 1200 kg.
AKA: L-9. Status: Active. Thrust: 27.40 kN (6,160 lbf). Gross mass: 12,500 kg (27,500 lb). Unfuelled mass: 2,700 kg (5,900 lb). Specific impulse: 324 s. Burn time: 1,100 s. Height: 3.36 m (11.02 ft). Diameter: 3.96 m (12.99 ft). Span: 5.46 m (17.91 ft).

Cost $ : 6.000 million.
http://www.astronautix.com/a/ariane5-2.html

Note that it's actual dry mass without the extra equipment used for its Ariane 5 function would only be 1,200 kg. Given also its high Isp, it could serve as a full stage descent stage, i.e., no separate crash landing for the stage.

Also, nice is its low cost.

  Bob Clark

Do you know of examples of those low chamber pressure solid thrusters at Pc = 100 psia? For my upper stage of the all-solid launcher, I want to use one of the size of the Cesaroni motor used on the Up Aerospace suborbital rocket I mentioned in post #13 with a 412 pound fueled weight and 0.8 propellant fraction. This puts is it at about 80 pound empty weight and 320 pound propellant weight. If I got the chamber pressure smaller by a factor of 10, then I could reduce that empty weight to only 8 pounds, and the payload would increase by 72 pounds.

But this means I want to keep the same propellant weight, and use the same size case but only thinner wall thickness by a factor of 10. But I don't know if this is possible for a solid-fueled motor. For a liquid I could just reduce the propellant burn rate to reduce the chamber pressure, i.e., less propellant burn at any time, less pressurized gases produced at any time, therefore lower chamber pressure.

But decreasing the chamber pressure is not so easy to do with a solid. It won't work to just increase the case volume because that eliminates the advantage you wanted to get of having a smaller empty, i.e., case weight. You can't reduce the propellant amount either to get a larger volume for the central open section because you want the same size stage.

Is there some method to get the same amount of propellant to just burn at a slower rate? Another possible way it could work would be by having the propellant compressed to an extreme degree to only 1/10th it's usually volume. Don't know if that is technically feasible.

  Bob Clark

Speaking of SSTO's, there are some solid stage motors with high mass ratio, in the range of 20 to 1. If this was kerolox or hydrolox with their high maximum vacuum Isp's this would be enough for a SSTO. Unfortunately, for solids their vacuum Isp's are typically only in the range of 285 s.

But I was surprised running the engine simulation program RPA, http://propulsion-analysis.com/index.htm, that the APCP solids could get vacuum Isp's in the range of 325 s and above by using nozzles with extremely high area ratios, in the range of hundreds to 1. Then running the launch performance calculator, http://silverbirdastronautics.com/LVperform.html, I found they could be SSTO.

Here's one solid motor with the high mass ratio:

Star 48
Thiokol solid rocket engine. Used in Delta 3900; Conestoga; PAM-S; PAM-D. Total flown included in total for Star-48-8. Total impulse 575,682 kgf-sec. Motor propellant mass fraction 0.946. First flight 1982.
AKA: PAM-D;PAM-S;TE-M-711-3. Status: Out of Production. Number: 125 . Thrust: 67.20 kN (15,107 lbf). Gross mass: 2,114 kg (4,660 lb). Unfuelled mass: 114 kg (251 lb). Specific impulse: 287 s. Burn time: 88 s. Height: 2.04 m (6.69 ft). Diameter: 1.24 m (4.06 ft).
Chamber Pressure: 40.00 bar. Area Ratio: 49.
http://www.astronautix.com/s/star48.html

Here's how the input screen looks on the launch performance estimator for this motor:

Note there some quirks of this program you need to be aware of if you use it. First, always use the vacuum values for the Isp's and thrust numbers, since the program already takes into account the diminution at sea level. Second, always set the "Restartable Upper Stage" option to "No", rather than the default "Yes", otherwise the payload will be reduced. This is true even for a supposed SSTO.
Third, always set the launch inclination to match the launch site latitude, otherwise the payload will be reduced. This is related to the fact that changing the orbital plane involves a delta-v cost. So for the Cape Canaveral launch site, the launch inclination should be set to 28.5 degrees.

Now, here's the result:

i.e., 0 payload is the estimation.

Now, for the case where the vacuum Isp can be raised to 325 s by using the aerospike or other altitude compensation methods the results are then:

i.e., 33 kg is the payload estimation.

  Bob Clark

What do you calculate the Isp would be for a 1,000 psi and 2,000 psi engine with an area ratio of, say, 500 to 1? As you say that would make a regular bell nozzle impractically large, well outside the motor's diameter. But with an aerospike it would stay inside the motors diameter. Surprisingly, after knowing about the aerospike for over 50 years there still have been no high altitude tests of the aerospike where the ambient pressure would be near vacuum. The only flight tests have only been at about 30,000 ft. where the pressure is still 1/3 sea level.

BTW, what do you use as the solid motors throat area when it is not circular but cross-shaped? Also it would seem to me the throat area should change as the solid fuel burns away.

  Bob Clark

Some amateur groups are working on hybrid engines. However, they are not as safe as portrayed, as the accident at Scaled Composites shows. Also they are not as well understood as solids and liquids, as the 10 year development cycle to get the hybrid engine on SpaceShipTwo to work properly shows.

   Bob Clark

This page from the Cesaroni web page says their high power rocket motors use ammonium perchlorate composite propellant (APCP):

http://www.pro38.com/regulations.php

They sell mostly to amateurs so perhaps they don't make the propellant with the optimal requirements to optimize performance.

  BTW, I was reading some articles of applications of nanotubes, among the now numerous applications of them I found some refs that found adding nanotubes to fuels speeds up the combustion speed to an extraordinary degree:

Highly energetic compositions based on functionalized carbon nanomaterials.
Qi-Long Yan a, Michael Gozin *a, Feng-Qi Zhao b, Adva Cohen a and Si-Ping Pang c
DOI: 10.1039/C5NR07855E (Review Article) Nanoscale, 2016, 8, 4799-4851
http://pubs.rsc.org/en/content/articleh … c5nr07855e [free full text]

Enhanced Energy Release from Homogeneous Carbon Nanotube-Energetic Material Composite.
Authors: Um, Jo-Eun; Yeo, Taehan; Choi, Wonjoon; Chae, Joo Seung; Kim, Hyoun Soo; Kim, Woo-Jae
Source: Science of Advanced Materials, Volume 8, Number 1, January 2016, pp. 164-170(7)
Publisher: American Scientific Publishers
http://www.ingentaconnect.com/content/a … cation/pdf [abstract only]

I wonder if this could increase the burn rate for the solid propellants even at relatively low pressures.

  Bob Clark