|
Appendix
C: Lander Mass Calculation and Launch Vehicle Overview
Lander Mass Calculation
The initial burnout mass
(Mbo) is assumed as follows:
The factor k is usually
a ratio of payload mass to spacecraft dry mass, but in this paragraph,
k is assumed to be a ratio of payload mass to initial burnout
mass for simplification. The current calculations assume k=25%.
The propellant mass is calculated as follows (Larson, W. and
Pranke, L., 1999):
Mass of engines, tanks,
and lines that depends on propellant mass is usually scaled as
follows:
With this information,
the propellant mass can be iterated again:
The total mass of the
lander (including payload) is given by:
For propellant the lander
uses the storable NTO and MMH liquid combination. The tables
below summarize the characteristic of mass of the lander. The
ratio of payload mass to spacecraft dry mass is finally 20%,
which is consistent with (Wertz, J. and Larson, W., 1992). This
value provides a conservative evaluation of the lander mass.
Table C-1: Payload mass
|
Payload |
Symbol |
Unit |
Value |
|
Rover mass |
MR |
kg |
10.0 |
|
The number of rovers |
NR |
- |
10 |
|
Media Vehicle |
MV |
kg |
15.0 |
|
Total mass |
Mtotal |
kg |
115.0 |
Table C-2:
Spacecraft Mass and sizing, intermediate results
|
Lander |
Symbol |
Unit |
Lunar Descent |
|
Initial payload ratio |
k |
- |
25% |
|
Initial burnout mass |
Mbo |
kg |
460.0 |
|
Delta v |
DV |
m/sec |
3050 |
|
Oxidizer/Fuel |
|
|
NTO/MMH |
|
Isp |
|
sec |
310 |
|
Propellant mass-1 (First iteration) |
Mprop-1 |
kg |
795.4 |
|
Engines, tanks, and lines |
0.15Mprop- |
kg |
119.3 |
|
Propellant mass-2 (Second iteration) |
Mprop-2 |
kg |
1001.6 |
|
Total stage mass |
Mtotal |
kg |
1580.9 |
Table C-3:
Mass and size calculation results
|
Results |
Unit |
Value |
|
Payload |
kg |
120 |
|
Other subsystems except propulsion subsystem |
kg |
350 |
|
Engines, tanks and feeding lines |
kg |
120 |
|
Spacecraft dry mass |
kg |
580 |
|
The ratio of payload mass to spacecraft
dry mass |
- |
20% |
|
Propellant |
ton |
1.0 |
|
Total lander mass (dry mass + propellants
mass) |
ton |
1.6 |
Table C-4:
Launcher Capability
|
Launcher |
P/L GTO (kg) |
Launch Site |
Orbit Incl (° ) |
Est. Launch Cost (Mln. US$) |
Reliability |
Avg. Cost / Reliability (Mln.
US$) |
Remarks |
|
Long March 3A |
2500 |
Xichan |
28 |
35-40 |
3/3
33/36 |
39.13 |
no Earth escape yet |
|
Delta II 7925 |
1842 |
Cape Canaveral |
28 |
45-50 |
54/56 |
49.26 |
could provide injection |
|
Atlas I |
2255 |
Cape Canaveral |
28 |
65-75 |
8/11 |
96.25 |
could provide injection |
|
Atlas II |
2810 |
Cape Canaveral |
28 |
75-85 |
5/5
19/21 |
76.19 |
could provide injection |
|
A40 |
2050 |
Kourou |
7 |
45-60 |
6/6
72/75 |
53.57 |
out of Service in 2002? |
|
A42P |
2840 |
Kourou |
7 |
60-75 |
10/11
72/75 |
72.23 |
out of Service in 2002? |
|
GSLV |
2500 |
Sriharikota |
18 |
? |
0/0 |
? |
not yet operational |
|
Zenit 3 |
5180 |
Tyuratam |
X |
50-70 |
26/32 |
73.85 |
|
|
Sealaunch |
5100 |
Ocean |
? |
60-90 |
1/1 |
75.00 |
no Earth escape yet |
|
Proton D1e |
4100 Lunar Transfer |
Tyuratam |
Lunar Transfer |
50-70 |
154/167 |
65.06 |
would provide injection |
|
H II |
4000 |
Tanegashima |
28.5 |
150-190 |
1/1 |
170.00 |
no Earth escape yet |
|
Molniya |
1600 Lunar Transfer |
Baikonur (B) / Plesetsk (P) |
Lunar Transfer |
12-39 |
46/71 (B)
183/190 (P)
229/261 |
39.36 26.48 29.06 |
would provide injection |
(Isakowitz, Steven J., 1991)
(Mark Wade, 1999)
NEXT >
[Home] [ISU] |
