Merge pull request #36 from alopezrivera/delfi-c3-properties-correction

Mass, reference area and initial state of Delfi-C3 corrected in all examples
tudat-team · Nov 3, 2023 · 88a602a · 88a602a
2 parents 7ff85a9 + 33acba3
commit 88a602a
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diff --git a/estimation/covariance_estimated_parameters.ipynb b/estimation/covariance_estimated_parameters.ipynb
@@ -46,6 +46,7 @@
     "from tudatpy import constants\n",
     "from tudatpy.interface import spice\n",
     "from tudatpy import numerical_simulation\n",
+    "from tudatpy.numerical_simulation import environment\n",
     "from tudatpy.numerical_simulation import environment_setup\n",
     "from tudatpy.numerical_simulation import propagation_setup\n",
     "from tudatpy.numerical_simulation import estimation, estimation_setup\n",
@@ -136,10 +137,10 @@
    "source": [
     "# Create vehicle objects.\n",
     "bodies.create_empty_body(\"Delfi-C3\")\n",
-    "bodies.get(\"Delfi-C3\").mass = 400.0\n",
+    "bodies.get(\"Delfi-C3\").mass = 2.2\n",
     "\n",
     "# Create aerodynamic coefficient interface settings\n",
-    "reference_area = 4.0\n",
+    "reference_area = (4*0.3*0.1+2*0.1*0.1)/4  # Average projection area of a 3U CubeSat\n",
     "drag_coefficient = 1.2\n",
     "aero_coefficient_settings = environment_setup.aerodynamic_coefficients.constant(\n",
     "    reference_area, [drag_coefficient, 0.0, 0.0]\n",
@@ -148,7 +149,7 @@
     "environment_setup.add_aerodynamic_coefficient_interface(bodies, \"Delfi-C3\", aero_coefficient_settings)\n",
     "\n",
     "# Create radiation pressure settings\n",
-    "reference_area_radiation = 4.0\n",
+    "reference_area_radiation = (4*0.3*0.1+2*0.1*0.1)/4  # Average projection area of a 3U CubeSat\n",
     "radiation_pressure_coefficient = 1.2\n",
     "occulting_bodies = [\"Earth\"]\n",
     "radiation_pressure_settings = environment_setup.radiation_pressure.cannonball(\n",
@@ -242,7 +243,7 @@
     "### Define the initial state\n",
     "Realise that the initial state of the spacecraft always has to be provided as a cartesian state - i.e. in the form of a list with the first three elements representing the initial position, and the three remaining elements representing the initial velocity.\n",
     "\n",
-    "Within this example, we will make use of the `keplerian_to_cartesian_elementwise()` function  - included in the `element_conversion` module - enabling us to convert an initial state from Keplerian elements to a 6x1 cartesian vector."
+    "Within this example, we will retrieve the initial state of Delfi-C3 using its Two-Line-Elements (TLE) the date of its launch (April the 28th, 2008). The TLE strings are obtained from [space-track.org](https://www.space-track.org)."
    ]
   },
   {
@@ -252,17 +253,13 @@
    "metadata": {},
    "outputs": [],
    "source": [
-    "# Set the initial state of the vehicle\n",
-    "earth_gravitational_parameter = bodies.get(\"Earth\").gravitational_parameter\n",
-    "initial_state = element_conversion.keplerian_to_cartesian_elementwise(\n",
-    "    gravitational_parameter=earth_gravitational_parameter,\n",
-    "    semi_major_axis=7500.0E3,\n",
-    "    eccentricity=0.1,\n",
-    "    inclination=np.deg2rad(85.3),\n",
-    "    argument_of_periapsis=np.deg2rad(235.7),\n",
-    "    longitude_of_ascending_node=np.deg2rad(23.4),\n",
-    "    true_anomaly=np.deg2rad(139.87)\n",
-    ")"
+    "# Retrieve the initial state of Delfi-C3 using Two-Line-Elements (TLEs)\n",
+    "delfi_tle = environment.Tle(\n",
+    "    \"1 32789U 07021G   08119.60740078 -.00000054  00000-0  00000+0 0  9999\",\n",
+    "    \"2 32789 098.0082 179.6267 0015321 307.2977 051.0656 14.81417433    68\"\n",
+    ")\n",
+    "delfi_ephemeris = environment.TleEphemeris( \"Earth\", \"J2000\", delfi_tle, False )\n",
+    "initial_state = delfi_ephemeris.cartesian_state( simulation_start_epoch )"
    ]
   },
   {

diff --git a/estimation/covariance_estimated_parameters.py b/estimation/covariance_estimated_parameters.py
@@ -32,6 +32,7 @@
 from tudatpy import constants
 from tudatpy.interface import spice
 from tudatpy import numerical_simulation
+from tudatpy.numerical_simulation import environment
 from tudatpy.numerical_simulation import environment_setup
 from tudatpy.numerical_simulation import propagation_setup
 from tudatpy.numerical_simulation import estimation, estimation_setup
@@ -92,10 +93,10 @@
 
 # Create vehicle objects.
 bodies.create_empty_body("Delfi-C3")
-bodies.get("Delfi-C3").mass = 400.0
+bodies.get("Delfi-C3").mass = 2.2
 
 # Create aerodynamic coefficient interface settings
-reference_area = 4.0
+reference_area = (4*0.3*0.1+2*0.1*0.1)/4  # Average projection area of a 3U CubeSat
 drag_coefficient = 1.2
 aero_coefficient_settings = environment_setup.aerodynamic_coefficients.constant(
     reference_area, [drag_coefficient, 0.0, 0.0]
@@ -172,20 +173,16 @@
 """
 Realise that the initial state of the spacecraft always has to be provided as a cartesian state - i.e. in the form of a list with the first three elements representing the initial position, and the three remaining elements representing the initial velocity.
 
-Within this example, we will make use of the `keplerian_to_cartesian_elementwise()` function  - included in the `element_conversion` module - enabling us to convert an initial state from Keplerian elements to a 6x1 cartesian vector.
+Within this example, we will retrieve the initial state of Delfi-C3 using its Two-Line-Elements (TLE) the date of its launch (April the 28th, 2008). The TLE strings are obtained from [space-track.org](https://www.space-track.org).
 """
 
-# Set the initial state of the vehicle
-earth_gravitational_parameter = bodies.get("Earth").gravitational_parameter
-initial_state = element_conversion.keplerian_to_cartesian_elementwise(
-    gravitational_parameter=earth_gravitational_parameter,
-    semi_major_axis=7500.0E3,
-    eccentricity=0.1,
-    inclination=np.deg2rad(85.3),
-    argument_of_periapsis=np.deg2rad(235.7),
-    longitude_of_ascending_node=np.deg2rad(23.4),
-    true_anomaly=np.deg2rad(139.87)
+# Retrieve the initial state of Delfi-C3 using Two-Line-Elements (TLEs)
+delfi_tle = environment.Tle(
+    "1 32789U 07021G   08119.60740078 -.00000054  00000-0  00000+0 0  9999",
+    "2 32789 098.0082 179.6267 0015321 307.2977 051.0656 14.81417433    68"
 )
+delfi_ephemeris = environment.TleEphemeris( "Earth", "J2000", delfi_tle, False )
+initial_state = delfi_ephemeris.cartesian_state( simulation_start_epoch )
 
 
 ### Create the integrator settings

diff --git a/estimation/estimation_dynamical_models.ipynb b/estimation/estimation_dynamical_models.ipynb
@@ -102,7 +102,7 @@
     "bodies.get(\"MEX\").mass = 1000.0\n",
     "\n",
     "# Create radiation pressure settings\n",
-    "reference_area_radiation = 4.0\n",
+    "reference_area_radiation = (4*0.3*0.1+2*0.1*0.1)/4  # Average projection area of a 3U CubeSat\n",
     "radiation_pressure_coefficient = 1.2\n",
     "occulting_bodies = [\"Mars\"]\n",
     "radiation_pressure_settings = environment_setup.radiation_pressure.cannonball(\n",

diff --git a/estimation/full_estimation_example.ipynb b/estimation/full_estimation_example.ipynb
@@ -48,6 +48,7 @@
     "from tudatpy import constants\n",
     "from tudatpy.interface import spice\n",
     "from tudatpy import numerical_simulation\n",
+    "from tudatpy.numerical_simulation import environment\n",
     "from tudatpy.numerical_simulation import environment_setup\n",
     "from tudatpy.numerical_simulation import propagation_setup\n",
     "from tudatpy.numerical_simulation import estimation, estimation_setup\n",
@@ -138,10 +139,10 @@
    "source": [
     "# Create vehicle objects.\n",
     "bodies.create_empty_body(\"Delfi-C3\")\n",
-    "bodies.get(\"Delfi-C3\").mass = 400.0\n",
+    "bodies.get(\"Delfi-C3\").mass = 2.2\n",
     "\n",
     "# Create aerodynamic coefficient interface settings\n",
-    "reference_area = 4.0\n",
+    "reference_area = (4*0.3*0.1+2*0.1*0.1)/4  # Average projection area of a 3U CubeSat\n",
     "drag_coefficient = 1.2\n",
     "aero_coefficient_settings = environment_setup.aerodynamic_coefficients.constant(\n",
     "    reference_area, [drag_coefficient, 0.0, 0.0]\n",
@@ -150,7 +151,7 @@
     "environment_setup.add_aerodynamic_coefficient_interface(bodies, \"Delfi-C3\", aero_coefficient_settings)\n",
     "\n",
     "# Create radiation pressure settings\n",
-    "reference_area_radiation = 4.0\n",
+    "reference_area = (4*0.3*0.1+2*0.1*0.1)/4  # Average projection area of a 3U CubeSat\n",
     "radiation_pressure_coefficient = 1.2\n",
     "occulting_bodies = [\"Earth\"]\n",
     "radiation_pressure_settings = environment_setup.radiation_pressure.cannonball(\n",
@@ -244,7 +245,7 @@
     "### Define the initial state\n",
     "Realise that the initial state of the spacecraft always has to be provided as a cartesian state - i.e. in the form of a list with the first three elements representing the initial position, and the three remaining elements representing the initial velocity.\n",
     "\n",
-    "Within this example, we will make use of the `keplerian_to_cartesian_elementwise()` function  - included in the `element_conversion` module - enabling us to convert an initial state from Keplerian elements to a 6x1 cartesian vector."
+    "Within this example, we will retrieve the initial state of Delfi-C3 using its Two-Line-Elements (TLE) the date of its launch (April the 28th, 2008). The TLE strings are obtained from [space-track.org](https://www.space-track.org)."
    ]
   },
   {
@@ -254,17 +255,13 @@
    "metadata": {},
    "outputs": [],
    "source": [
-    "# Set the initial state of the vehicle\n",
-    "earth_gravitational_parameter = bodies.get(\"Earth\").gravitational_parameter\n",
-    "initial_state = element_conversion.keplerian_to_cartesian_elementwise(\n",
-    "    gravitational_parameter=earth_gravitational_parameter,\n",
-    "    semi_major_axis=7500.0E3,\n",
-    "    eccentricity=0.1,\n",
-    "    inclination=np.deg2rad(85.3),\n",
-    "    argument_of_periapsis=np.deg2rad(235.7),\n",
-    "    longitude_of_ascending_node=np.deg2rad(23.4),\n",
-    "    true_anomaly=np.deg2rad(139.87)\n",
-    ")"
+    "# Retrieve the initial state of Delfi-C3 using Two-Line-Elements (TLEs)\n",
+    "delfi_tle = environment.Tle(\n",
+    "    \"1 32789U 07021G   08119.60740078 -.00000054  00000-0  00000+0 0  9999\",\n",
+    "    \"2 32789 098.0082 179.6267 0015321 307.2977 051.0656 14.81417433    68\"\n",
+    ")\n",
+    "delfi_ephemeris = environment.TleEphemeris( \"Earth\", \"J2000\", delfi_tle, False )\n",
+    "initial_state = delfi_ephemeris.cartesian_state( simulation_start_epoch )"
    ]
   },
   {

diff --git a/estimation/full_estimation_example.py b/estimation/full_estimation_example.py
@@ -32,6 +32,7 @@
 from tudatpy import constants
 from tudatpy.interface import spice
 from tudatpy import numerical_simulation
+from tudatpy.numerical_simulation import environment
 from tudatpy.numerical_simulation import environment_setup
 from tudatpy.numerical_simulation import propagation_setup
 from tudatpy.numerical_simulation import estimation, estimation_setup
@@ -92,10 +93,10 @@
 
 # Create vehicle objects.
 bodies.create_empty_body("Delfi-C3")
-bodies.get("Delfi-C3").mass = 400.0
+bodies.get("Delfi-C3").mass = 2.2
 
 # Create aerodynamic coefficient interface settings
-reference_area = 4.0
+reference_area = (4*0.3*0.1+2*0.1*0.1)/4  # Average projection area of a 3U CubeSat
 drag_coefficient = 1.2
 aero_coefficient_settings = environment_setup.aerodynamic_coefficients.constant(
     reference_area, [drag_coefficient, 0.0, 0.0]
@@ -172,20 +173,16 @@
 """
 Realise that the initial state of the spacecraft always has to be provided as a cartesian state - i.e. in the form of a list with the first three elements representing the initial position, and the three remaining elements representing the initial velocity.
 
-Within this example, we will make use of the `keplerian_to_cartesian_elementwise()` function  - included in the `element_conversion` module - enabling us to convert an initial state from Keplerian elements to a 6x1 cartesian vector.
+Within this example, we will retrieve the initial state of Delfi-C3 using its Two-Line-Elements (TLE) the date of its launch (April the 28th, 2008). The TLE strings are obtained from [space-track.org](https://www.space-track.org).
 """
 
-# Set the initial state of the vehicle
-earth_gravitational_parameter = bodies.get("Earth").gravitational_parameter
-initial_state = element_conversion.keplerian_to_cartesian_elementwise(
-    gravitational_parameter=earth_gravitational_parameter,
-    semi_major_axis=7500.0E3,
-    eccentricity=0.1,
-    inclination=np.deg2rad(85.3),
-    argument_of_periapsis=np.deg2rad(235.7),
-    longitude_of_ascending_node=np.deg2rad(23.4),
-    true_anomaly=np.deg2rad(139.87)
+# Retrieve the initial state of Delfi-C3 using Two-Line-Elements (TLEs)
+delfi_tle = environment.Tle(
+    "1 32789U 07021G   08119.60740078 -.00000054  00000-0  00000+0 0  9999",
+    "2 32789 098.0082 179.6267 0015321 307.2977 051.0656 14.81417433    68"
 )
+delfi_ephemeris = environment.TleEphemeris( "Earth", "J2000", delfi_tle, False )
+initial_state = delfi_ephemeris.cartesian_state( simulation_start_epoch )
 
 
 ### Create the integrator settings

diff --git a/propagation/keplerian_satellite_orbit.ipynb b/propagation/keplerian_satellite_orbit.ipynb
@@ -225,12 +225,12 @@
     "earth_gravitational_parameter = bodies.get(\"Earth\").gravitational_parameter\n",
     "initial_state = element_conversion.keplerian_to_cartesian_elementwise(\n",
     "    gravitational_parameter=earth_gravitational_parameter,\n",
-    "    semi_major_axis=7500.0e3,\n",
-    "    eccentricity=0.1,\n",
-    "    inclination=np.deg2rad(85.3),\n",
-    "    argument_of_periapsis=np.deg2rad(235.7),\n",
-    "    longitude_of_ascending_node=np.deg2rad(23.4),\n",
-    "    true_anomaly=np.deg2rad(139.87),\n",
+    "    semi_major_axis=6.99276221e+06,\n",
+    "    eccentricity=4.03294322e-03,\n",
+    "    inclination=1.71065169e+00,\n",
+    "    argument_of_periapsis=1.31226971e+00,\n",
+    "    longitude_of_ascending_node=3.82958313e-01,\n",
+    "    true_anomaly=3.07018490e+00,\n",
     ")"
    ]
   },

diff --git a/propagation/keplerian_satellite_orbit.py b/propagation/keplerian_satellite_orbit.py
@@ -3,7 +3,6 @@
 Copyright (c) 2010-2022, Delft University of Technology. All rights reserved. This file is part of the Tudat. Redistribution and use in source and binary forms, with or without modification, are permitted exclusively under the terms of the Modified BSD license. You should have received a copy of the license with this file. If not, please or visit: http://tudat.tudelft.nl/LICENSE.
 """
 
-
 ## Context
 """
 This example demonstrates the basic propagation of a (quasi-massless) body under the influence of a central point-mass attractor. It therefore resembles the classic two-body problem.
@@ -148,12 +147,12 @@
 earth_gravitational_parameter = bodies.get("Earth").gravitational_parameter
 initial_state = element_conversion.keplerian_to_cartesian_elementwise(
     gravitational_parameter=earth_gravitational_parameter,
-    semi_major_axis=7500.0e3,
-    eccentricity=0.1,
-    inclination=np.deg2rad(85.3),
-    argument_of_periapsis=np.deg2rad(235.7),
-    longitude_of_ascending_node=np.deg2rad(23.4),
-    true_anomaly=np.deg2rad(139.87),
+    semi_major_axis=6.99276221e+06,
+    eccentricity=4.03294322e-03,
+    inclination=1.71065169e+00,
+    argument_of_periapsis=1.31226971e+00,
+    longitude_of_ascending_node=3.82958313e-01,
+    true_anomaly=3.07018490e+00,
 )
 
 ### Create the propagator settings

diff --git a/propagation/linear_sensitivity_analysis.ipynb b/propagation/linear_sensitivity_analysis.ipynb
@@ -46,6 +46,7 @@
     "# Load tudatpy modules\n",
     "from tudatpy.interface import spice\n",
     "from tudatpy import numerical_simulation\n",
+    "from tudatpy.numerical_simulation import environment\n",
     "from tudatpy.numerical_simulation import environment_setup, propagation_setup, estimation_setup\n",
     "from tudatpy.astro import element_conversion\n",
     "from tudatpy import constants\n",
@@ -143,7 +144,7 @@
     "bodies.get(\"Delfi-C3\").mass = 400.0\n",
     "\n",
     "# Create aerodynamic coefficient interface settings\n",
-    "reference_area = 4.0\n",
+    "reference_area = (4*0.3*0.1+2*0.1*0.1)/4  # Average projection area of a 3U CubeSat\n",
     "drag_coefficient = 1.2\n",
     "aero_coefficient_settings = environment_setup.aerodynamic_coefficients.constant(\n",
     "    reference_area, [drag_coefficient, 0.0, 0.0]\n",
@@ -153,7 +154,7 @@
     "            bodies, \"Delfi-C3\", aero_coefficient_settings)\n",
     "\n",
     "# Create radiation pressure settings\n",
-    "reference_area_radiation = 4.0\n",
+    "reference_area_radiation = (4*0.3*0.1+2*0.1*0.1)/4  # Average projection area of a 3U CubeSat\n",
     "radiation_pressure_coefficient = 1.2\n",
     "occulting_bodies = [\"Earth\"]\n",
     "radiation_pressure_settings = environment_setup.radiation_pressure.cannonball(\n",
@@ -256,7 +257,7 @@
     "\n",
     "This initial state always has to be provided as a cartesian state, in the form of a list with the first three elements reprensenting the initial position, and the three remaining elements representing the initial velocity.\n",
     "\n",
-    "In this case, let's make use of the `keplerian_to_cartesian_elementwise()` function that is included in the `element_conversion` module, so that the initial state can be input as Keplerian elements, and then converted in Cartesian elements."
+    "Within this example, we will retrieve the initial state of Delfi-C3 using its Two-Line-Elements (TLE) the date of its launch (April the 28th, 2008). The TLE strings are obtained from [space-track.org](https://www.space-track.org)."
    ]
   },
   {
@@ -266,17 +267,13 @@
    "metadata": {},
    "outputs": [],
    "source": [
-    "# Set the initial state of the vehicle\n",
-    "earth_gravitational_parameter = bodies.get(\"Earth\").gravitational_parameter\n",
-    "initial_state = element_conversion.keplerian_to_cartesian_elementwise(\n",
-    "    gravitational_parameter=earth_gravitational_parameter,\n",
-    "    semi_major_axis=7500.0E3,\n",
-    "    eccentricity=0.1,\n",
-    "    inclination=np.deg2rad(85.3),\n",
-    "    argument_of_periapsis=np.deg2rad(235.7),\n",
-    "    longitude_of_ascending_node=np.deg2rad(23.4),\n",
-    "    true_anomaly=np.deg2rad(139.87)\n",
-    ")"
+    "# Retrieve the initial state of Delfi-C3 using Two-Line-Elements (TLEs)\n",
+    "delfi_tle = environment.Tle(\n",
+    "    \"1 32789U 07021G   08119.60740078 -.00000054  00000-0  00000+0 0  9999\",\n",
+    "    \"2 32789 098.0082 179.6267 0015321 307.2977 051.0656 14.81417433    68\"\n",
+    ")\n",
+    "delfi_ephemeris = environment.TleEphemeris( \"Earth\", \"J2000\", delfi_tle, False )\n",
+    "initial_state = delfi_ephemeris.cartesian_state( simulation_start_epoch )"
    ]
   },
   {