DOI: 10.5281/zenodo.17580678
title: CARIOQA Quantum Pathfinder Mission Phase A orbit and accelerometer data

description: This dataset accompanies the paper CARIOQA Quantum Pathfinder Mission for space weather research submitted to the Geodesy for a Changing Environment - Proceedings of the 2025 IAG Scientific Assembly.

The CARIOQA Quantum Pathfinder Mission aims at demonstrating a quantum technology-based accelerometer in space as a precursor for a later deployment onboard a satellite gravimetry mission. A dedicated satellite will be launched for the Pathfinder Mission in the early 2030s to raise the technology level of the required technologies for applications on operational missions.

The Phase A study to investigate the feasibility of this mission has concluded and the project will continue into Phase B in October of 2025. We present studies from the Phase A on the relationship between available models of the atmospheric conditions in low Earth orbits and the instrument and satellite design as well as the impact on the development of requirements to fulfil the mission objectives. In addition to the demonstration of the functionality of the quantum accelerometer, the Pathfinder Mission will also provide accelerometer measurements in low Earth orbits for the expected mission lifetime of three years. As a scientific objective of the mission, this dataset will be used to derive parameters like thermosphere density or atmospheric crosswinds.
description: Definitions

For details we refer to the paper to be published in the IAG Proceedings.

Orbits

The following sun-synchronous orbits are considered in this study



CIRC: circular orbit with 500km altitude

ELLIP: elliptical orbit between 400km and 700km altitude.


Models

The software VENQS was used to simulate the orbit of the satellite considering its 3D geometry and models of the satellites environment



Atmospheric drag: NRLMSISE-00 (F10.7 = 115, Ap = 13)

Solar Radiation Pressure: Point-like sun; eclipse included

Gravity field: GGM05C


Frames

The satellite is controlled to be nadir pointing. The following frames are defined:



Satelite body frame



x-axis is pointing along track

y-axis is normal to orbital plane pointing towards sun

z-axis is pointing radial outward



Sensor Frame



x-axis is pointing along track

y-axis is normal to orbital plane pointing away from sun

z-axis nadir pointing




 
description: The data is provided in csv file format with file names according to the following convention:

YYYY-MM-DD_VENQS_data_fileX_ORBIT_Fall_24h



YYYY-MM-DD: date of creation of dataset

fileX: filetype 1 or 2 (see following tables)

ORBIT: orbit scenario ELLIP or CIRC


The simulated data is provided for 24h starting on 22.09.2030 12:00 with a 4s sampling intervall.

Data formats

Filetype 1

This file includes satellite orbit information.




Column
Title
Description


1
MJD [d]
Time stamp as Modified Julian Date


2
pos. x [m] 
x component of satellite position in ECI (i) frame (ICRF)


3
pos. y [m] 
y component of satellite position in ECI (i) frame (ICRF)


4
pos. z [m] 
z component of satellite position in ECI (i) frame (ICRF)


5
vel. x [m/s]
x component of satellite velocity in ECI (i) frame (ICRF)


6
vel. y [m/s]
y component of satellite velocity in ECI (i) frame (ICRF)


7
vel. z [m/s]
z component of satellite velocity in ECI (i) frame (ICRF)


8
qs
Scalar component of quaternion of satellite altitude




9

q1
1st imaginary part




10

q2
2nd imaginary part




11

q3
3rd imaginary part




12

omega x [rad/s]
x-component of angular velocity in the satellite body frame (b)




13

omega y [rad/s]
y-component of angular velocity in the satellite body frame (b)




14

omega z [rad/s]
z-component of angular velocity in the satellite body frame (b)




The provided quaternions transform a vector in ECI frame (i) into satellite body frame (b)

Filetype 2

This filetype includes accelerations and gravity gradients data.




Column
Title
Description


1
MJD [d]
Time stamp as Modified Julian Date


2
Drag Acc-x [m/s^2]
x-component of satellite acceleration due to atmospheric drag


3
Drag Acc-y [m/s^2]
y-component of satellite acceleration due to atmospheric drag


4
Drag Acc-z [m/s^2]
z-component of satellite acceleration due to atmospheric drag


5
SRP Acc-x [m/s^2]
x-component of satellite acceleration due to solar radiation pressure


6
SRP Acc-y [m/s^2]
y-component of satellite acceleration due to solar radiation pressure


7
SRP Acc-z [m/s^2]
z-component of satellite acceleration due to solar radiation pressure


8
GG-matrix 1,1 [1/s^2]
G11


9
GG-matrix 1,2 [1/s^2]
G12


10
GG-matrix 1,3 [1/s^2]
G13


11
GG-matrix 2,1 [1/s^2]
G21


12
GG-matrix 2,2 [1/s^2]
G22


13
GG-matrix 2,3 [1/s^2]
G23


14
GG-matrix 3,1 [1/s^2]
G31


15
GG-matrix 3,2 [1/s^2]
G32


16
GG-matrix 3,3 [1/s^2]
G33




The accelerations are provided in the satellite sensor frame.

The gravity gradient matrix (GG) is defined at the centre of mass of the satellite in satellite body frame (b) with the components Gxy (x:row y:column).