# EPFLx: EE585x Space Mission Design and Operations

EPFLx: EE585x Space Mission Design and Operations
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Welcome to the course Space Mission Design and Operations!

Space exploration is truly fascinating. From the Sputnik mission to the assembly and exploitation of the International Space Station or the Apollo program, we are just beginning to uncover the many mysteries of our universe!

This course may help you decide if you really want to become an aerospace engineer or, why not, an astronaut.

This course builds on university-level mechanics to introduce and illustrate orbital dynamics as they are applied in the design of space missions. You will learn from the experiences of Claude Nicollier, one of the first ESA astronauts, specifically, his role in the maintenance of the Hubble Space Telescope with NASA.

The course will start with a brief history (and mainly human) of space exploration , an introduction to main space missions before reviewing the fundamental laws of newtonian mechanics and conservation laws. It will present some basic concepts on space environment, robotic and human spacecraft and will introduce operational aspects of such vehicles, including maneuvers and propulsion.

The list of concepts that will be explained during the course is also available here

Learners will gain knowledge of the challenges related to the use of the space environment as a scientific and commercial platform.

We hope that you will enjoy the course!

Concepts explained in this course:
C1 - Newton's laws of motion

First law: In the absence of a force, a body either is at rest or moves in a straight line with constant speed.

Second law: A body experiencing a force will be subject to an acceleration such that the force is equal to the product of the mass of the body and the acceleration.

Third law: Whenever a first body exerts a force on a second body, the second body exerts a force equal in magnitude and opposite in direction on the first body.

This concept is explained in the following video: 1.2.1.

C2 - Inertial Frame

The inertial frame is a frame with respect to which the laws of Newton are valid.

The direction of the axes of an inertial frame can be imagined as being fixed with respect to distant stars.

The center of the inertial frame, which is an orthogonal coordinate system, will depend on the application.

This concept is explained in the following video: 1.2.1.

C3 - Newton's law of gravitation

Newton's law of gravitation states that any two bodies in the universe attract each other with a force that is directly proportional to the product of their masses and inversely proportional to the square of the distance between them.

This concept is explained in the following video: 1.2.2.

C4 - Laws of motion of a solid body and in rotation

This concept is based on the Newton’s second law of motion applied for a rotation: the torque is equal to the time derivative of the angular momentum.

This concept is explained in the following video: 1.2.2.

C5 - Conservation Laws

Conservation of momentum in an isolated system (absence of forces).

Conservation of angular momentum in an isolated system (absence of torques).

Conservation of mechanical energy, potential and kinetic, in an isolated system, in a conservative force field (for instance a gravitational force field in the absence of dissipative forces).

This concept is explained in the following video: 1.2.3.

C6 - Structure and composition oF the Earth's atmosphere

Chemical composition: 78 % of nitrogen, 21 % of oxygen, 1 % of argon, < 0.1 % others.

Layers: Troposphere, stratosphere, mesosphere, thermosphere. Tropopause: limit of the troposphere, 9 to 17 km altitude, depending on the latitude. Stratopause: limit of stratosphere, around 50 km altitude.

The limit of the atmosphere and the beginning of space is considered to be at 100 km altitude.

This concept is explained in the following video: 1.3.1.

C7 - Microgravity

Microgravity is the term used to characterize the very low acceleration level encountered inside a spacecraft in LEO (typically

g at 300 km altitude).

This concept is explained in the following video: 1.3.1.

C8 - Transparency of the atmosphere

The atmosphere is opaque to electromagnetic radiation except in the visible part of the spectrum and in the radio part between about 3 cm to 10 m wavelength. It has limited transparency in the infrared, between 1 and 10 microns wavelength.

This concept is explained in the following video: 1.3.2.

C9 - Light Effects

Airglow: emission of light by the atmosphere due to the photoionization of oxygen atoms and de-excitation which produces a luminescence, mainly due to the oxygen, somewhat also to nitrogen and the radical OH.

This concept is explained in the following video: 1.3.2.

C10 - Magnetic Field

Close to the Earth’s surface, the geomagnetic field is essentially a bipolar field slightly offset from the center of the Earth. The expression of the amplitude of the magnetic field is a function of the distance to the center of the Earth and the magnetic latitude.

Radiation or Van Allen belts: High energy protons and electrons trapped in two regions of the magnetosphere, Protons and electrons in the inner RB, electrons only in the outer RB.

The Earth's magnetic field is significantly distorted away from the surface of Earth toward the Sun (5, up to 10 Earth radii) and in the anti-Sun direction due to solar wind, made of charged particles flowing in all directions from the Sun.

This concept is explained in the following videos: 1.4.1.

C11 - Northern/Southern lights

Northern and Southern Lights are produced from the excitation of nitrogen and oxygen atoms, by the electrons flowing from the solar wind.

This concept is explained in the following videos: 1.4.1.

C12 -Solar cycle - influence on the Earth’s atmosphere

A solar cycle is a period of approximately 11 years, based on the sunspot number, which is changing over time with an increase of sunspot during 5 years until the solar maxima followed by 6 years of decrease until the solar minima.

A better determination of the phase of the solar cycle is the solar radio flux. The solar radiation flux at a wavelength of 10.7 cm varies along the solar cycle and is often used as a measure of the level of the solar activity along the cycle.

The Sun has also a significant effect on the thickness of the atmosphere; respectively on the density of the Earth’s atmosphere at a given altitude.

This concept is explained in the following videos: 1.4.2.

C13 - Sun activity

The Sun is an active star. Its surface is granular with prominences: flares and coronal mass ejections.

A solar prominence is a large, bright feature extending outward from the Sun’s surface, often in a loop shape. While the corona consists of extremely hot ionised gases, which do not emit much visible light, prominences contain much cooler plasma, similar in composition to that of the chromosphere.

This concept is explained in the following video: 1.4.3.

C14 - Particle flux

Following Coronal Mass Ejections or CMEs, there is an enormous amount of charged particles that flow through the Solar System all the way to the Earth and beyond.

This concept is explained in the following video: 1.5.1.

C15 - South Atlantic Anomaly

The South Atlantic Anomaly or SAA is an area where the Earth’s inner Van Allen radiation belt comes closest to the Earth’s surface.

This concept is explained in the following video: 1.5.1.

C16 - Solar radiation, outside the Earth’s atmosphere and on the surface

Cosmic particles of galactic origin: high-energy charged particles, usually protons, electrons and fully ionized nuclei of light elements.

Radiation are expressed in RAD - Radiation Absorbed Dose - which is the amount of energy absorbed or REM (Roentgen Equivalent Man); or Sievert (Sv = 100 REM).

Astronauts have to stay way below 100 REM, or one Sievert, which is the vomiting effective threshold. Medical problems start over 100 REM and become significant above 500 REM.

This concept is explained in the following video: 1.5.1.

The total irradiance of the Sun is the total energy that it sends every unit of time or every second on a square meter surface outside of the Earth's atmosphere, in the form of electromagnetic radiation.

The solar constant is a conventional measure of the mean total solar irradiance at a distance of one Astronomical Unit.

Solar irradiance in the Solar system decreases with the distance to the Sun.

This concept is explained in the following video: 1.5.2.

C18 - Albedo

Albedo is the diffuse reflectivity or reflecting power of a surface measured from zero for no reflecting power of a perfectly black surface, to 1 for perfect reflection.

This concept is explained in the following video: 1.5.3.

C19 - Earth’s energy budget

The solar constant is approximately 1368 watts per square meter. This value has been slowly decreasing over the past few years and in April 2015, the value was 1361 watts per square meter.

About 30% of this radiation is reflected back in space by the atmosphere (6%), clouds (20%) and the Earth's surface (4%). The rest is absorbed by land and oceans and re-radiated out in space in the form of infrared radiation.

From the Earth, there are two sources of radiation: the albedo, which is the solar spectrum reflected from the high layers of the Earth atmosphere, and the infrared radiation from the Earth itself.

This concept is explained in the following video: 1.5.3.

C20 - Radiation balance for a spacecraft on orbit around the Earth

If we consider a spacecraft in the vicinity of the Earth, the dominant source of radiation it receives comes from the Sun. 30% of the radiation of the Sun impacting the Earth is coming back towards space (Albedo). The other source of radiation is the Earth infrared radiation, which is a rather dim source.

The total absorbed radiation by the spacecraft is equal to the emitted radiation in a balanced situation.

The Stefan-Boltzmann law gives the value of the emitted radiation per unit surface in the case of a black body.

The spacecraft is not a black body and its emitted radiation depends of its total surface and its IR emissivity.

The equation expressing the radiation balance can be used to determine the average temperature of a spacecraft, knowing its emissivity and absorptivity.

This concept is explained in the following video: 1.5.4.

C21 - Orbit decay - ballistic coefficient

The orbit decay is the reduction in the altitude of a satellite's orbit. The major cause is the drag of the Earth’s atmosphere, which is more important during the solar maxima at a given altitude. Positive Feedback effect: The more the orbit decays, the lower the altitude, the faster the decay.

The drag equation expresses the force experienced by an object moving through a fluid or a gas at a velocity V.

The ballistic coefficient is the measure of the resistance to orbit decay caused by atmospheric drag and is inversely proportional to the drag coefficient and the frontal surface of the spacecraft.

This concept is explained in the following video: 1.6.1.

C22 - Space debris

Space debris is an object in space that is no longer functional and useful. There is a huge amount of space debris in LEO and GEO causing high risk of collision with active spacecraft.

Several rules have been put in place to regulate satellites allocation in certain areas and the end of missions: when a communication satellite in the geostationary orbit are is longer functional, it shall be moved to a graveyard orbit. Measures are in place to limit satellites lifetime in Low Earth Orbit to 25 years.

Two break-up: Iridium-Cosmos in 2009 and Fengyun-1C in 2007 caused a very high density of space debris around 800 km altitude.

This concept is explained in the following video: 1.6.2.

C23 - asteroid collision threat

Several asteroids collisions happened on Earth but also in other planets in the Solar System.

A major event was the Cretaceous-Paleogene extinction, 65 million years ago, large scale extinction of most of life on Earth, including dinosaurs from the impact of large meteorite Earth, probably in the region of the Yucatan peninsula, Mexico.

Different techniques to improve early detection of dangerous asteroids, and to modify their trajectories to avoid collision with earth are currently being looked at.

This concept is explained in the following video: 1.6.3.

C24 - Gravitational Profile of the Earth

Inside the Earth, considered homogeneous, the gravitational acceleration varies linearly from zero at the center to 9.81 meters per second square at the surface. From the surface on, it varies as 1/r2, r being the distance to the center of the Earth.

This concept is explained in the following video: 2.2.1.

C25 - Gravitational well, Work and Depth

A gravitational well is a conceptual model of the gravitational field surrounding a body in space.

The work necessary to lift a unit mass from the surface of a spherical object such as Earth, to infinity, is equal to the work necessary to lift the same mass from the surface of the sphere over a distance equal to the object’s radius, with a constant force equal to the force at the surface.

The object’s radius is the depth of the object’s gravitational well. The profile of the gravitational well is in 1/r

This concept is explained in the following videos: 2.2.1. and 2.2.2.

C26 - Normalized gravitational well

The depth of gravitational well of any spherical objects in the solar system or elsewhere, is always normalized to the gravitational acceleration of the Earth for comparison purposes.

This concept is explained in the following video: 2.2.2.

C27 - Escape Velocity

Escape velocity from the Earth’s surface is the velocity at which a spacecraft has to leave this surface in order to go to infinity with a zero velocity. The escape velocity is independent of the direction of the initial impulse (as long as escape really takes place).

For Earth, the escape velocity from the surface is equal to 11.2 km/second. For higher altitudes, it varies as

.

This concept is explained in the following video: 2.2.3.

C28 - Circular Velocity

The circular velocity is the velocity of a spacecraft on a circular orbit around an object such as the Earth. For a spacecraft in orbit around the Earth, its circular velocity is equal to

.

This concept is explained in the following video: 2.2.3

C29 - GRavitational well in term of transfer velocity

The transfer velocity, for a given planet, is the velocity that has to be added to the planet’s circular velocity for a transfer to infinity from this location in the Sun’s gravitational well, i.e. as if to leave the solar system.

This concept is explained in the following video: 2.2.3.

C30 - Two-Body Problem

The two-body problem is to determine the motion of the two bodies that interact only with each other.

This concept is explained in the following video: 2.3.1.

C31 - Kepler's Laws

First law: The orbit of every planet is an ellipse with the Sun at one of the two foci.

Second law: A line is joining a planet and the Sun sweeps out equal areas during equal intervals of time.

Third law: The square of the orbital period of a planet is proportional to the cube of the semi-major axis of the orbit.

This concept is explained in the following video: 2.3.1.

C32 - Generalisation of the first Kepler's Law

The first Kepler’s law can be generalized in the case of a two-body problem: the orbit of the small body versus the large body is, generally speaking, a conic, i.e. an ellipse, a circle, a parabola or hyperbola.

This concept is explained in the following video: 2.3.1.

C33 - Important elliptical parameters: Periapsis, apoapsis, eccentricity, true anomaly

Periapsis and apoapsis are general terms. Periastris and apoastris are sometimes used for a star as central body. If the Earth is the central body, we talk about perigee and apogee; if it is the Sun, perihelion and aphelion.

If the eccentricity is equal to 0 the orbit is circular, if the distance of the focus from the center of the orbit increases, the eccentricity becomes closer to 1. If the eccentricity is equal to 1, the orbit is parabolic.

The True anomaly is the angle between the direction of the periapsis from the central body to the radius vector to the spacecraft or the planet.

This concept is explained in the following video: 2.4.1.

C34 - Mechanical Energy of the orbital motion

The total mechanical energy of a body in a gravitational field is the sum of its kinetic energy and its potential energy.

In the limit case of a very elongated ellipse, the total mechanical energy of the orbital motion is close to zero.

In the limit case of a parabolic orbit, the total mechanical energy is equal to zero.

If V < escape velocity, which is the case for a closed orbit, elliptical or circular, the total mechanical energy is negative.

If the orbit is hyperbolic (velocity > 0 at infinity), the total mechanical energy is positive.

This concept is explained in the following video: 2.4.1.

C35 - Orbital Velocity

The orbital velocity for an elliptical orbit is

The orbital velocity for a circular orbit is

This concept is explained in the following video: 2.4.1.

C36 - Flight Path Angle

The flight path angle is the angle between the direction of the velocity vector and the perpendicular to the radius vector at the point where the spacecraft is.

The variation of the flight path angle on an elliptical orbit as a function of position:

This concept is explained in the following video: 2.4.1.

C37 - Circular and elliptical orbits – fundamental relationshis

Several formulas express fundamental relationships between elliptical parameters.

Some of the most used formulas are the expression of the mean motion and the Kepler’s equation.

The mean motion is the average angular speed required for a body to complete one orbit, it is expressed in radians per second.

E, the eccentric anomaly, is an angular parameter that defines the position of a body that is moving along an elliptic Kepler orbit.

The Kepler’s equation is a transcendental equation that cannot be solved for E but expresses the time evolution of E, the eccentric anomaly.

This concept is explained in the following video: 2.4.2.

C38 - Reference Frames

The geographic coordinate system is used for describing positions on the surface of planet Earth.

The geocentric-inertial coordinate system is an orthogonal reference frame centered at the center of the Earth. The plane of reference is the plane of the Equator, where the direction of X is the direction of the vernal equinox for a given year, nowadays chosen to be the year 2000. Y is also in the equator, and Z in the rotation axis of the Earth pointing to the North

The heliocentric-inertial coordinate system has the same directions of axes as the geocentric-inertial coordinate system, but it is centered at the center of the Sun.

This concept is explained in the following video: 2.5.1.

C39 - AXial Precession

Axial precession is the displacement of the rotational axis of an astronomical body.

The Earth is not a perfect sphere, but has an equatorial bulge, and the gravitational force from the Sun and the Moon, on a non-spherical body, causes the precession. The Earth goes through one such complete precessional cycle in about 26000 years.

This concept is explained in the following video: 2.5.1.

C40 - Classical orbital parameters and spacecraft state vector

There are 6 orbital parameters used to describe an orbit: the eccentricity, the semi-major axis, the inclination, the longitude or Right Ascension of the Ascending Node (RAAN, in the plane of reference), the argument of periapsis (in the orbital plane), the time of periapsis transit.

With the addition of the time t, we can have a determination of the exact position of the celestial body or satellite.

The spacecraft’s state vector is functionally equivalent to the six orbital parameters plus the time t.

This concept is explained in the following video: 2.5.2.

C41 - Mean solar day and sidereal day

The sidereal day, is the time it takes for the Earth to make one full rotation with respect to the stars.

The mean solar day is the time it takes for the Earth to make one full rotation with respect to the Sun.

The duration of the mean solar day is 24 hours, but the duration of the sidereal day, is 23h 56’ 04’’.

This concept is explained in the following video: 2.5.2.

C42 - Calendars

Julian day is used in the Julian date (JD) system of time measurement for scientific use by the astronomy community, presenting the interval of time in days and fractions of a day since January 1st, 4713 BC Greenwich noon. Julian date is recomended for astronomical use by the International Astronomical Union.

This concept is explained in the following video: 2.5.2.

C43 -Maneuvers in Orbit

A maneuver on orbit is the modification of the orbit of a satellite by adding a velocity (ΔV) on the initial orbit.

This concept is explained in the following video: 3.2.1.

C44 - Hohmann Transfer

The Hohmann transfer is a transfer between one circular orbit to another circular orbit around the same central body.

This concept is explained in the following video: 3.2.1.

C45 - Hohmann transfer for small increments in speed or distance to the central body

This Hohmann transfer is often used for rendezvous at Low Earth Orbit.

For LEO:

This concept is explained in the following video: 3.2.2.

C46 - Orbital change plane

The orbital plane change or orbital inclination change is an orbital maneuver aimed at changing the inclination of an orbiting body's orbit.

This concept is explained in the following video: 3.2.3.

C47 - Geosynchronous and Geostationary orbits

If the rotation period of a satellite is equal to a sidereal day, the satellite will rotate around the Earth at the same pace as the Earth rotates around itself: this is called geosynchronous orbit.

A Geostationary orbit is a geosynchronous circular orbit on the equatorial plane (e = 0 and i = 0).

This concept is explained in the following video: 3.3.1.

C48 - strategy for reaching such an orbit, and changing spacecraft longitude

To reach the geostationary orbit, a Hohmann transfer is necessary followed by a plane change of the orbit from the initial plane to the Equatorial condition. The apogee shall be at the geostationary altitude, 36,000 km above the Earth's surface.

There are two options to reach the geostationary orbit:

a plane change and then an acceleration to the circular condition at that orbit.
a combined maneuver of the plane change and acceleration to the circular condition

This concept is explained in the following video: 3.3.1.

C49 -Geometry of a launch from the Earth’s surface

A satellite has a certain inclination versus the Equator which is equal to the latitude of the launch site. Typically, seven degrees for a launch from French Guiana, in Kourou, the European launch site, 28.5 degrees if the launch takes place from Kennedy Space Center, Florida.

This concept is explained in the following video: 3.3.1.

C50 - Nodal regression (or progression) of an orbit around the Earth

One consequence of the Earth’s equatorial bulge is the nodal regression, for orbits with an inclination of more than 0 and less than 90, the line of nodes is moving to the west.

For inclinations larger than 90 degrees, the line of node will be drifting to the east: this is the nodal progression.

This concept is explained in the following video: 3.3.2.

C51 - Sun synchronous orbits

Sun-synchronous orbit is an orbit that keeps the same orientation versus the Sun as the Earth is going around the Sun in a full year.

This concept is explained in the following video: 3.3.2.

C52 -Restricted three-body problem

The restricted three-body problem is applied to two relatively large bodies and a smaller body, the spacecraft. The two main bodies are on circular orbits around the center of mass of the system. The mass of the third body (satellite) is very small compared to the mass of the two main bodies. The third body is in an orbit contained in the plane of the orbits of the two main bodies.

This concept is explained in the following video: 3.3.3.

C53 -Lagrange points

The Lagrange points are five positions where the small body keeps the same position with respect to the other two bodies which are revolving around each other.

This concept is explained in the following video: 3.3.3.

C54 -Catch up rate for nearby orbits

In general, for circular and close elliptical orbits:
with

For circular orbits:
with

This concept is explained in the following video: 3.4.1.

C55 - Maneuvers and Burns

As orbital altitude increases, orbital period increases and orbital velocity decreases. Velocity is greatest at perigee and least at apogee.

Posigrade burns increase altitude 180° from the burn point. Retrograde burns decrease altitude 180° from the burn point. Radial burns shift the semi-major axis without significantly altering other orbital parameters.

Plane change must be performed at junction of two orbits: nodal crossing.

This concept is explained in the following video: 3.4.2.

C56 -Rendezvous in Earth's Orbit - General Strategy

The rendezvous is the action of bringing together two spacecraft. Most of the time there is a spacecraft on the ground, called chaser, which is active and another spacecraft on orbit, the target, passive.

In the initial conditions, the orbit inclination is larger than the latitude of the launch site for the chaser.

The final conditions are chaser and target at the same location in space with the same vectorial velocity.

Example of the the Space Shuttle and HST, Space Shuttle and ISS, HTV and ISS, ATV and ISS.

This concept is explained in the following videos: 3.4.3. and 3.5.1.

C57 -Phase Angle and Phasing Rate

Phase angle is the angle between the chaser and target, measured from the center of the Earth. Phasing rate or catch-up rate is the rate at which phase angle changes. Phasing rate is a function of differential altitude.

This concept is explained in the following videos: 3.4.4.

C58 -Rendezvous profile

The profile of the orbit of the chaser versus a target is represented in a one-dimensional plane which is a plane of the orbit of the target, at least a plane of the orbit of the target at the end of the rendezvous

This concept is explained in the following videos: 3.4.4.

C59 - RENDEZVOUS Control

Rendezvous sensors are used to update the relative state vector of a chaser versus target using sensor data, to control the rotation and translation and to perform manually the end portion of the rendezvous.

3 sensors on the Space Shuttle : Star Tracker (S TRK), Rendezvous Radar (RR) and Crew Optical Alignment Sight (COAS).

Rotation hand controller (RHC): 3 degrees of freedom rotation: pitch, roll and yaw.

Translation hand controller (THC): translation on X or Z axis.

This concept is explained in the following videos: 3.4.5.

C60 - Relative motion of the chaser vs. the target in a rendezvous

A posigrade burn of the Shuttle with respect to the Station implies a higher and slower orbit with a longer period.

A retrograde burn brings the Shuttle on an orbit with lower altitude behind the Station.

A radial burn is a burn perpendicular to the velocity vector. The amplitude of the velocity vector, the energy and the period will be unchanged.

This concept is explained in the following videos: 3.4.6.

C61 -Case of the ATV

The ATV brings resupply equipment, fuel, water, payloads and food for the crew. It can also reboost the ISS, firing thrusters with the proper attitude, to the nominal altitude.

The ATV performs automatic rendezvous and automatic re-entry after de-docking. Its re-entry is destructive.

This concept is explained in the following video: 3.5.1.

C62 - Astronomical unit

Astronomical unit is the average distance between the Sun and the Earth

1 AU = 149.5978707 x 106 km

This concept is explained in the following video: 4.2.1.

C63 -Interplanetary Trajectories

In order to plan for and execute a mission to another planet, we consider the Sun, the planet of departure (the Earth), the planet of destination and the spacecraft.

When the spacecraft leaves the sphere of influence of the Earth, it occurs to be on a heliocentric elliptical trajectory, from the departure orbit, which is the orbit of the Earth, until the destination orbit which is a semi-circular orbit, either inside the orbit of the Earth or outside for Mars, Jupiter or Saturn, etc.

This concept is explained in the following video: 4.2.1.

C64 -Patched conics approximation

The spacecraft is on orbit around the Earth, on a parking orbit, and is accelerated to leave the area of the Earth, the sphere of influence.

As long as the spacecraft is within this sphere of influence, its motion with respect to the Earth is a two-body problem with the Earth as a central body and the spacecraft having a very small mass compared to the mass of the Earth.

At the end its heliocentric elliptic trajectory, in the vicinity of the destination planet, the spacecraft enters the sphere of influence of the destination planet (two-body problem).

This concept is explained in the following video: 4.2.1.

C65 -SPHERE OF INFLUENCE

Sphere around each planet inside which the motion of a spacecraft is considered to be two-body Keplerian.

The radius of the sphere of influence RS has been determined by Laplace as:

R is the average distance between Sun and the planet.

This concept is explained in the following video: 4.2.1.

C66 -Symbol Convention for Position and Velocity

P for planet, S for Spacecraft, Capital letters for heliocentric movements, small letters for planetocentric movements.

The heliocentric movement of the spacecraft is equal to the planetocentric motion of the spacecraft added to the heliocentric movement of the planet when the spacecraft leaves the sphere of influence.

If the spacecraft enters a sphere of influence, the planetocentric movement of the planet is equal to the difference between the heliocentric movement of the spacecraft and the planets.

This concept is explained in the following video: 4.2.2.

C67 -Departure from a planet

A journey to a destination in the solar system always starts with a parking orbit. Then the orbital velocity in increased to reach the departure velocity.

The departure velocity amplitude influences the orbit of the spacecraft: if the departure velocity is smaller than the escape velocity, the spacecraft will be on an elliptical orbit, if it is equal, the spacecraft will be on a parabola, if it is higher, the spacecraft will be on a hyperbolic orbit.

This concept is explained in the following video: 4.2.2. and 4.2.5.

C68 -Hyperbolic excess velocity & reaching the Sphere of Influence

On a hyperbolic orbit, at a large distance from the Earth, the spacecraft comes to a certain constant velocity

.

This concept is explained in the following videos: 4.2.2. and 4.2.3.

C69 -Vis-viva equation - elliptical and Hyperbolic velocities

Ellipse :

Hyperbola

This concept is explained in the following video: 4.2.3. and 4.2.5.

C70 -Arrival at a planet

The spacecraft is on a heliocentric trajectory, elliptical orbit around the Sun, until it is close enough to the destination planet. Then only the motion of the spacecraft with respect to the destination planet is considered as hyperbolic trajectory inside the sphere of influence of the destination planet.

Flight controllers in the control room control the motion of the spacecraft and select the value of d∞, the impact parameter, and θ in order to accomplish either a flyby or landing.

ΔVi indicates how much the spacecraft needs to brake in km per second in order to insert itself in an elliptical orbit around the destination planet.

This concept is explained in the following videos: 4.2.4. and 4.2.5.

C71 -Aerobraking, Aerocapture, Aeroentry

The idea is to generate changes of the velocity.

It could be done by using the atmosphere of a planet (Mars, Venus) or the satellite of a planet like Titan, which will minimize the use of propellant or by thrusters.

Aerocapture : Transfers the spacecraft from a hyperbolic approach trajectory to an elliptical orbit around the target planet. Further loss of energy will occur at every subsequent crossing of the periapsis.

Aerobraking : Transfers the spacecraft from an initial elliptical orbit to a less energetic (i.e. lower apoapsis) elliptical orbit. Involves small ΔV.

Aeroentry : Transfers the spacecraft from either a hyperbolic, parabolic or elliptical approach orbit to the planet surface

This concept is explained in the following video: 4.3.1.

C72 -Gravity assist or slingshot maneuver

When a spacecraft on an orbit around the Sun approaches another planet and comes in close proximity to it, the gravity of that planet takes over, pulling the spacecraft in and altering its speed.

The amount by which the spacecraft speeds up or slows down is determined by the geometry of the approach, passing behind or in front of the planet.

When the spacecraft leaves the sphere of influence of the planet, it again follows an orbit around the Sun, but a different one from before, either on course for the original target or heading for another fly-by.

This concept is explained in the following video: 4.3.2

C73 -Tsiolkovsky Equation

ΔV = change of velocity; ve = exhaust velocity of the gas in the propulsion system; mi, mf = initial and final mass

Valid in free space – Gravitational field-induced and drag-induced ΔVs will be added to the propulsion-induced ΔVs.

This concept is explained in the following video: 4.4.1.

C74 -Specific Impulse

Isp, the specific impulse, is a measure of the propulsion system efficiency, it is its thrust (kg-force) divided by the mass flow of propellants (fuel and oxidizer, kg/s).

This concept is explained in the following video: 4.4.1.

C75 -Chemical propulsion

Chemical propulsion is used for all of the stages until orbit insertion. Two main types of chemical propulsion systems: monopropellant or bipropellant.

The fuel and the oxidizers are being fed in the combustion chamber in liquid form or the solid propellant contains the oxidizer and the fuel in solid form

An igniter system is needed if there is no hypergolic propellant and oxidizer. A hot gas is produced under high pressure, expanded in a nozzle in order to increase the velocity of the exhaust.

This concept is explained in the following video: 4.4.1. and 4.4.2

C76 -Nuclear Propulsion

Heating of the propellant with a nuclear reactor with the control of the flux of neutrons with a reflector.

The velocity that can be reached is higher than with a chemical rocket.

This concept is explained in the following video: 4.4.3.

C77 -Electric or ion Propulsion

Ionization of propellants and acceleration with the electric field.

Higher velocities than with a liquid-fueled or solid propellant rocket engine can be reached.

There is high efficiency or very high ejection velocity, but relatively low thrust, of the order of a fraction of a newton.

It can be used for propulsion in space, but not for leaving the Earth's surface and t bring a spacecraft to orbital conditions.

This concept is explained in the following video: 4.4.3

C78 -Strategies for ascent into orbit, and reentry in the Earth’s atmosphere

The idea is to shape the ascent trajectory to minimize gravity and drag losses.

Orbit insertion consists in bringing a spacecraft to a desired stable orbit after a launch from the Earth surface.

Direct insertion into orbit: initial launch at the vertical, powered ascent, using the propulsion system of either one, two or three stages until orbit insertion.

For Earth launch, the ascent trajectory shall be lofted because of the atmosphere. On a planet with thinner atmosphere like Mars, loft is less necessary.

Braking maneuver such as deorbit burn are realized for the re-entry of the Space Shuttle.

● Entry requirements and constraints applicable for any re-entry vehicle, which does not have a destructive re-entry: Deceleration: Human limit is about 12g’s for short duration.

● Heating: Must withstand both total heat load and peak heating rate.

● Accuracy of landing or impact: Function primarily of trajectory and vehicle design.

● Size of the entry corridor: The size of the corridor depends on three constraints (deceleration, heating and accuracy).

This concept is explained in the following video: 4.5.1.

C79 -Attitude Measurement and control system (AMCS)

Man-made spacecraft or natural objects like asteroid or comet nucleus, are very slowly rotating subject to possibly gravity gradient forces, if located in the vicinity of a large body, to sun radiation, solar wind, atmospheric effect of a near planet or magnetic effect.

AMCS consists in measuring and maintaining, or changing in a controlled manner, the orientation of a coordinate system attached to the spacecraft with respect to an inertial or any other reference system. Attitude within a specified deadband.

This concept is explained in the following video: 5.2.1.

C80 - Euler sequence

Rotations around axis X, Y, and Z are non commutative. Normally Euler sequence: Yaw Pitch Roll. For the Space Shuttle, ISS, including associated robotic system: Pitch Yaw Roll.

This concept is explained in the following video: 5.2.1.

C81 -types of attitude control

Gravity gradient, Magnetic torquers, Spinning spacecraft, Momentum devices: reaction wheels or Control Momentum Gyros, Thrusters. Passive and active methods.

This concept is explained in the following video: 5.2.2.

An elongated object will take an orientation in orbit around the Earth such that its long axis will be along the local vertical, with some swinging motion, or attitude changes around the local vertical gravity gradients.

This concept is explained in the following video: 5.2.2.

C83 -Magnetic torquers

A magnetic torquer is an elongated bar with a wire coil wrapped around it and an external protection.

A current through the coil will produce a magnetic field which will try to align itself along the geomagnetic field with a torque expressed by:

Torques on the spacecraft induced by variations of the wheel’s rotational speed. Requires thrusters or magnetic torque generators to get out of wheels saturation.

Magnetic torquers are used if the orientation to space track doesn’t need to be extremely precise and as a system to desaturate momentum gyros (HST example).

This concept is explained in the following video: 5.2.2. and 5.2.3.

C84 -Stabilization by rotation

The inertial orientation is maintained by spinning the spacecraft with the sun vector perpendicular to the axis of rotation.

Advantages: cheap, propellant flow from tanks provided by inertial forces

Disadvantages: low accuracy in controlled attitude (0.3-1°), translations only possible along rotation axis, pointing of antennas and other devices impossible except in the direction of spin

This concept is explained in the following video: 5.2.2. and 5.2.3.

C85 -Three-axis stabilization with thrusters

Advantages: Any orientation possible, high pointing accuracy (better than 0.001°), efficient use of solar arrays.

Disadvantages: complexity and price, complex redundancy architecture, need to insure propellant supply from the tanks by other means than using the inertial forces.

This concept is explained in the following videos: 5.2.2., 5.2.3. and 5.2.4.

C86 - Reaction wheel, Control Momentum Gyros or CMGs

Constant angular velocity. Mounted on gimbals. A torque along the input axis produces a corresponding torque reaction along the output axis.

If the angular rotation speed of the reaction wheel is increased, the angular rotation speed of the spacecraft will increase in the opposite direction.

This concept is explained in the following videos: 5.2.3. and 5.2.5.

C87 - Electrical power generation and distribution, basic principles

Fuel cells, solar arrays, primary batteries, nuclear reactors, radioisotopic thermal generators.

Fuel cells convert chemical energy from reactants into electricity through a chemical reaction of positively charged hydrogen ions with oxygen (or other oxidizing agent). They require continuous source of reactants to sustain the chemical reaction. The by-products of this reaction are water and heat.

Rechargeable batteries: NiCd (Nickel Cadmium) currently replaced by Li-Ion (Lithium Ion) and NiH2 (Nickel Hydrogen)

A Radioisotope Thermoelectric Generator (RTG), uses the fact that radioactive materials (such as plutonium 238) generate heat as they decay into non-radioactive materials. The heat is converted into electricity by an array of thermocouples which then power the spacecraft.

This concept is explained in the following video: 5.3.1.

C88 -Solar Cells

An electrical energy storage system must be provided and must be sized as a function of the eclipse duration and electrical power requirements.

Upper layer: N doped semiconductor, Lower layer: P doped semiconductor, PN junction which, when irradiated by a photon flow, will displace electrons form N to P, holes from P to N from which voltage and current are resulting.

Cell types: Silicium or Gallium arsenide.

This concept is explained in the following video: 5.3.1.

C89 - Tethered satellite – concept and utilization (mechanical and electrodynamical)

Applications: Electrical power generation, Orbit transfers, Ionospheric studies, Variable gravity research, Space debris removal, Provision of artificial gravity for long journeys in the Solar System, Earth-Moon payload transfer, Space Elevator.

This concept is explained in the following videos: 5.3.2., 5.3.3. and 5.3.4.

C90 - Reliability and redundancy of space systems

Reliability R(t) is the probability that the system will not fail in the interval (0, t). MTTF = Mean Time To Failure, average time duration until first failure. MTBF = Mean Time Between Failures, average time duration between two consecutive failures.

This concept is explained in the following video: 5.4.1.

C91 -Space Shuttle conception and first tests

The conception of the Space Shuttle started in the 70s. The idea was to have a reusable spacecraft for access to low Earth orbit to launch commercial payloads, to use it for the needs of the Department of Defense and to do space exploration and utilization.

A fully reusable spacecraft would have induced quite high development cost for a relatively low operation cost, because everything was going to be reusable.

Due to budget limitations, NASA and the US government had to cut it off at the lower point of the horizontal axis, meaning less development costs but a higher cost per flights as the final configuration chosen was only partially reusable.

Approach and Landing Tests (ALT) in 1977, with different tail cone configuration and the Space Shuttle Enterprise.

This concept is explained in the following video: 6.2.1.

C92 -Space Shuttle final configuration

External tank, 2 solid rocket boosters, Orbiter with 3 main engines, protection tiles, large payload bay, cabin for up to 7 crewmembers. Protection tiles are able to resist high temperature during the re-entry. The heating was mainly in the nose of the Orbiter, the leading edge of the wings and the bottom of the Orbiter (black tiles). The white areas were also covered with thermal protection tiles, but not with the same capabilities that the black ones.

This concept is explained in the following video: 6.2.1.

C93 -Space Shuttle program

OFTs (Orbital Flight Tests): Tests of the system, STS-1, the first flight, in April, 1981, until STS-4 in 1982. Operational phase: STS-5, in 1982, through STS-135, in 2011.

John Young and Robert Crippen the two crewmembers of STS-1.

Major achievements: Hubble Program, Assembly of the ISS, 355 astronauts to go in space, 135 flights total.

Two major accidents in this program: Challenger in 1986, and Columbia in 2003

6 orbiters were used: Enterprise (Test vehicle), Columbia (28 missions), Challenger (10 missions), Atlantis (39 missions), Discovery (33 missions), Endeavour (25 missions).

This concept is explained in the following videos: 6.2.2. and 6.3.1.

C94 -Space Shuttle Habitable compartments

Forward portion: the flight deck, the mid-deck and the airlock.

Seats for ascent to orbit and re-entry: four in the flight deck, the commander, the pilot and two mission specialists and up to three in the mid deck.

Instrument panels, windows, lockers with personal items, tools and food, food preparation station to hydrate and heat the food, waste collection system.

The airlock was inhabited when there was space walking in order to go out and come back in the Space Shuttle.

This concept is explained in the following video: 6.2.3.

C95 -Space Shuttle from countdown to orbit

Access to the cabin with an elevator and a bridge giving access to the White Room, where the crew with the help of technicians was making the last preparations before getting into the Space Shuttle orbiter cabin.

Orange Launch and Entry Suit and helmet, there was a possibility of pressurizing the suit in case there was any problem with the cabin pressure during the ascent. There was also a possibility to technically explode the side hatch for rapid escape in case there was a launch abort and the crew had to escape rapidly before liftoff of the orbiter.

Solid Rocket Booster ignition at T-zero Initial vertical lift off then shortly after roll maneuver in order to come into the proper plane of the planned orbit which could be between 28.5 degrees (latitude of Kennedy Space Center) to 52 degrees (inclination of the International Space Station) Ascent until Solid Rocket Booster separation few seconds later, which fall down into the Pacific ocean and recovered for later flights. Then the ascent to orbit continues for 6.5 extra minutes, for 8.5 total minutes. And at the end of this ascent phase, there is an acceleration of the velocity of 20,000 km/hour or about Mach 76. Then, main engine cutoff and the Shuttle goes from 3 G to 0 G within a very short time - about a second and a half.

This concept is explained in the following video: 6.2.4.

C96 -Space Shuttle: From orbit to post landing

During very hot phase of the re-entry between about 80 km altitude and 50 km, there was about 10 minutes of generation of heat because of the friction of the orbiter coming in the atmosphere

The orbiter had a high angle of attack: about 40 degrees angle of attack to brake in the higher layers of the atmosphere and to enter the lower layers of the atmosphere at a lower speed to control the heating and the trajectory of the spaceship.

After touchdown there was a deployment of the brake chute, a slow slap-down, the nose of the orbiter was coming down until nose wheel touch down.

OPF: Orbiter Processing Facility - Maintenance during several months stacking in the vehicle assembly building. Complexity of hardware around the orbiter for the maintenance between flights, especially concerning the thermal protection system.

This concept is explained in the following video: 6.2.5.

C97 - Space Shuttle selected missions

Utilization of the Shuttle: Satellites deployment including Hubble Space Telescope, Return to Earth of broken satellites, Space science with mainly Spacelab - 22 missions from STS-9 in 1983 until 1998, Interventions on satellites or telescope, Repair work, Assembly of large structures (ISS).

Space Shuttle-MIR program. Roscosmos and NASA technical and human cooperation: design and engineering of mating system between the Shuttle and Mir, opportunity for US astronauts to work with Russian cosmonauts. The idea was to prepare the assembly and utilization of the International Space Station.

This concept is explained in the following video: 6.3.1.

C98 - Space Shuttle Challenger accident

STS 51L in January 1986: explosion of the Space Shuttle happened 73 seconds after lift off.

The cause of the accident was the faulty design of the field joint and the low temperature preventing the O-ring to perform its expected task.

The Rogers Commission recommended to add another O-ring and a heater system to maintain a relatively high temperature in the area of every field joint of the Solid Rocket Booster.

After the Challenger accident, it was decided to use the Shuttle only for research and development, science and not for commercial missions or Department of Defense missions.

The crew had to wear the Launch and Entry Suit to protect itself against possible failures during the ascent, like a loss of cabin pressure.

NASA took advantage on the basis of the recommendations of the Rogers Commission to improve the safety of the Space Shuttle and of the organization.

This concept is explained in the following video: 6.3.2.

C99 - Space Shuttle Columbia accident

Lift off in January 2003, the accident happened during the re-entry of the spaceship, two weeks later.

A piece of thermal protection of the external tank detached from the tank and hit the wing.

The main recommendations were to fix the problem of detachment of thermal protection from the external tank, and give a possibility to the crew to do inspections of the thermal protection system while on orbit to make sure that there were no fatal damage of this thermal protection, and to do repair by space walking with necessary tools and materials.

This concept is explained in the following video: 6.3.3

C100 - Space Shuttle Hubble servicing missions

5 servicing missions: SM1 in December 1993, SM2 – February 1997, SM3A – December 1999, SM3B – March 2002, SM4 – May 2009. All servicing missions served the purpose of repair or improvements of the telescope.

Intervention on orbit could be done in the instrument section or on the surface of the telescope (addition of thermal protection for magnetometers). The solar arrays were replaced several times

This concept is explained in the following video: 6.3.4.

C101 - International Space Station

Signature of ISS Agreements in 1998 by 15 countries: US, Russian Federation, Japan, Canada and 11 ESA member states.

The idea was to make available a world-class laboratory in space for life science, materials science, Earth observations, solar physics, technology development and support of long-duration spaceflight in the Solar System in the future.

Mating of the first two elements: Zarya Russian module and Unity node or Node 1, 1st US element, followed by another Russian module, Zvezda.

Two years after the beginning of the assembly of the Station, the first crew arrived by Soyuz.

The ISS has been permanently inhabited from that time on until now. Astronauts and cosmonauts stay five to six months missions on board.

This concept is explained in the following video: 6.4.1.

Since the Shuttle was retired in the summer of 2011, the only way to get to the International Space Station for astronauts and cosmonauts is using the Russian Soyuz, launched from Baikonur with a Soyuz rocket.

The "escape tower“, on top of the Soyuz capsule, allows the crew to escape in case there is any problem during launch. Rockets on this escape tower can be fired and take the Soyuz capsule, with the three crew members on board, to some distance and then come down with parachutes to the ground.

The nominal approach and docking of the Soyuz capsule to ISS is automatic, with various sensors, antennas and cameras that are used to guide and control the displacement of Soyuz in the vicinity of its target, ISS.

There are always several Soyuz attached to the ISS to rescue crewmembers in case of emergency and bring them back to Earth.

Re-entry: braking maneuver that initiate the descent, with fiery re-entry in the atmosphere, and large parachute opening. For the last portion of the descent, before touch down, there is a retrorocket firing at the last moment in order to reduce the shock of the contact of the descent module with the ground.

This concept is explained in the following video: 6.4.2.

3 commercial companies have been chosen by NASA to develop a manned system to and from ISS from US soil.

Space X is developing the Manned Dragon, human version of the Dragon capsule used to deliver freight to the ISS and return trash to Earth for 6 or 7 crewmembers

SpaceX has recently done an evaluation of the pad abort system of the manned dragon capsule, to be used in case there is a problem with a Falcon-9 rocket.

CST-100 capsule, developed by Boeing, will be able to take six to seven crew members to Station.

Sierra Nevada Corporation develops the Dream Chaser, a mini Shuttle that will be able from 2017-2018 to take crews to ISS.

This concept is explained in the following video: 6.4.2.

Shenzou 10, the last version of the ascent and descent vehicle, has three modules: on-orbit module; descent module, where the crew takes place for the ascent and the entry, and the service module with solar arrays on both sides.

China is developing right the successor to Tiangong 1 Space Station, Tiangong 2, and a new vehicle for ascent and re-entry, Shenzhou 11.

This concept is explained in the following video: 6.4.2.

C105 - Logistics supply to ISS

Progress – HTV – Dragon – ATV – Cygnus (2 versions)

NASA’s Commercial Crew and Cargo Program Office (C3PO): Commercial Orbital Transportion Services (COTS); Commercial Crew Development (CCDev) with Space X, Boeing and Sierra Nevada Corporation; Commercial Resupply Services (CRS): Phase 1: Space X and Orbital Sciences and Phase 2: Space X, Orbital ATK, Sierra Nevada Corporation

Delivery and return of 14-17 t of pressurized cargo in 4 or 5 trips per year. Delivery and disposal of 1.5-4 t of unpressuried cargo per year. Various ground support services for ISS resupply.

This concept is explained in the following video: 6.4.3.

C106 -Extravehicular Activities (EVA)

Extravehicular activity (EVA) is any activity done by an astronaut or cosmonaut outside a spacecraft beyond the Earth's appreciable atmosphere.

The first spacewalk was realised by Alexei Leonov in May 1965.

EVA can aborted because of problem threating the life of the astronaut such as a water leak inside helmet.

EVA have been used in Hubble servicing missions, ISS assembly, ISS repairs. Spacewalks have been mostly done from the Space Shuttle and the ISS.

This concept is explained in the following video: 7.2.1. and 7.2.3.

C107 - EVA space suits

Different type of Spacesuits exist depending of their utilization: High altitude pilot suit, Launch and Entry Suit, NASA Extravehicular Mobility Unit (EMU), NASA Manned Maneuvering Unit design (MMU), Orlan spacesuit.

MMU is a small spacecraft attached to the spacesuits that allowed astronauts to move around in translation rotation and to get toward the satellites without being attached to the robot arm, or without having a safety tether.

This concept is explained in the following video: 7.2.2.

C108 -Robotics arms for ISS and Space Shuttle

Shuttle Remote Manipulator System (SRMS) or Canadarm was used to rendezvous with Hubble and as a support to EVA.

Space Station Remote Manipulator System (SSRMS) or Canadarm 2.

Both RMS have been manufactured by Canada.

This concept is explained in the following video: 7.3.1.

C109 - Astronaut training

Centrifuge that creates the condition of ascent and re-entry, zero-g to experience microgravity, survival trainings, flight trainings, training of missions in simulators, mockups and pools, using virtual reality.

This concept is explained in the following video: 7.4.1.

C110 - Commercial space suborbital

A suborbital spaceflight is a flight in which the spacecraft reaches space (>100 km altitude in the case of the Earth), but its trajectory intersects the atmosphere or surface of the body from which it was launched, so that it does not even complete one orbit.

The first suborbital spaceflight was realized with Mercury-Redstone 3 and 4 in 1961.

The current actors of suborbital flights are: Virgin Galactic, XCOR Aerospace, Blue Origin and Airbus defense and Space.

This concept is explained in the following video: 7.5.1.

C110 - Commercial space suborbital

A suborbital spaceflight is a flight in which the spacecraft reaches space (>100 km altitude in the case of the Earth), but its trajectory intersects the atmosphere or surface of the body from which it was launched, so that it does not even complete one orbit.

The first suborbital spaceflight was realized with Mercury-Redstone 3 and 4 in 1961.

The current actors of suborbital flights are: Virgin Galactic, XCOR Aerospace, Blue Origin and Airbus defense and Space.

This concept is explained in the following video: 7.5.1.

c111 - Future of space utilization and exploration

Earth and atmosphere science and monitoring, Space applications (Galileo navigation system, GPS, etc…, communications), Space Solar Power concept, James Webb Space Telescope (JWST), Orion, rise of commercial space.

This concept is explained in the following video: 7.6.1.

C112 - Future human solar system exploration

Moon: ESA’s Moon village, exploiting the resources; Manned missions to asteroids; Mars: human settlement; Europa and Saturn moons.

This concept is explained in the following video: 7.6.2.

C113 - Orion and Space Launch System

Orion Multi-Purpose Crew Vehicle has been in development since 2005, in collaboration with ESA since 2013. It consists of a launch abort system tower, the crew module and the service module. Crew from 2 to 6 astronauts. Launched by Space Launch System (SLS), first crewed flight planned for 2021.

This concept is explained in the following video: 7.6.3.

Please note that the links to the videos do not work if the unit is not yet released.

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