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Analysis of Orbital Systems

  • Krafft A. Ehricke
Chapter

Abstract

The paper presents an analysis of orbital systems, consisting of the orbital establishment, its supply vehicles and their technique of operation. Satellite orbits are classified as permanent (stationary for more than 10 years) and as temporary. The lower altitude limit for permanent satellite orbits appears to be 450 to 500 miles. These orbits are occupied by observational satellites. An altitude range between 500 and 700 miles is found to be relatively most desirable for observational satellites. Three types of temporary orbits are defined, namely auxiliary orbits (120 to 150 miles altitude; duration of occupation a few hours at the most; purpose is payload transfer), orbits of departure of astronautical expeditions, particularly into the translunar space (350 to 400 miles altitude; duration of occupation about one year or less; purpose is assembly of inter-orbital vehicles), and orbits of arrival of astronautical expeditions (30,000 to 40,000 miles altitude for Venus or Mars expeditions; duration of occupation for a few days by returning expedition until being picked up by orbital vehicle from the earth). It is shown that the optimum satellite orbit of departure is as close to the earth as feasible and that the optimum orbit of return as well as the optimum satellite orbit at the target planet vary with the target planet. Therefore, with the exception of the orbit of arrival, all satellite orbits preferably lie below 700 miles altitude, that is, within the physical atmosphere of the earth.

Orbital supply systems are discussed, distinguishing between passenger-carrying and load-carrying orbital vehicles. The latter type is fully automatic. A large supply vehicle will be needed for periods of establishing orbital installations. For their maintenance a small version is anticipated. Desirable features and configurations for observational satellites and for orbital vehicles are discussed.

Nomenclature

CL

lift coefficient

F

thrust

Fc

apparent centrifugal force

i

orbital inclination with respect to the ecliptic

P

precession constant

R

distance from the sun

r

distance from the earth

Ys,opt

distance of optimum satellite orbit of departure with respect to a given interplanetary transfer ellipse

S

lifting area

Tpr

period of regression of the orbital nodes due to polar-precession

Tpr

period of regression with respect to the half of the globe being exposed to the sun

tr*

sidereal period of revolution in an orbit

v

velocity

W

vehicle weight

y

altitude

α

angle of attack

γ

or γ parameter of the terrestrial gravity field (6.254 · 104 n · mi3/sec2 = 3.98 · 105 km3/sec2)

γ

parameter of the solar gravity field (2.067 · 1010 n mi3/sec2 = = 1.36 · 1011km3/sec3)

δ

orbital inclination with respect to the equator

δ

displacement thickness of the boundary layer

θ

trajectory angle

λ

mean free molecular path

ϱ

air density

σ

ϱ y /ϱ 00

φ̄

angular velocity of satellite in orbit

φp

rate of regression

ω

angular velocity Subscripts

A

apogee or aphelion

c

circular

P

perigee or perihelion

s

satellite

y

referring to altitude y

00

surface value

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Copyright information

© Springer-Verlag Wien 1955

Authors and Affiliations

  • Krafft A. Ehricke
    • 1
  1. 1.Preliminary Design DepartmentBell Aircraft CorporationBuffaloUSA

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