Defining the Fidelity of Lifetime Calculations | Computing Lifetime | Reviewing Lifetime Results

Lifetime Tool

The Lifetime tool estimates the amount of time a satellite can be expected to remain in orbit before atmospheric drag and other perturbations cause it to decay.

The algorithms behind the Lifetime tool are designed to compute an estimate of duration of time that a satellite will remain in orbit when re-entry is not imminent. This tool is not designed to produce a precise time or location of re-entry. Full fidelity orbit propagation methods such as those available in HPOP and Astrogator are better suited for re-entry prediction.

To open the Lifetime tool, highlight the satellite in the Object Browser and select Lifetime... from the Satellite menu.

The Lifetime tool is available only for Satellites orbiting Earth.

Initial State

The initial state of the satellite specified on the satellite's Basic Orbit properties page will be interpreted by the Lifetime Tool in one of two ways, depending on the propagator selected for it. If the satellite's propagator is set to TwoBody, J2Perturbation, or J4Perturbation, the initial conditions are treated as mean orbital elements; if it is set to HPOP, SGP4, or LOP, these values represent the osculating state.

Options for Calculating Lifetime

The following table describes the options that you can use to further define the calculation that the tool will perform:

Option Description
Cd (Drag Coefficient) The satellite's drag coefficient, usually taken to be between 2.0 and 2.2.
Cr (Solar Radiation Pressure Coefficient) The satellite's SRP coefficient. A value of 0 indicates that the satellite is transparent to solar radiation; a value of 1 indicates that it is perfectly absorbing. A value of 4/3 means that it is flat, specularly reflecting.
Drag Area The mean cross-sectional area of the satellite perpendicular to its direction of travel.
Area Exposed to Sun The satellite's mean area projected perpendicular to the Sun's direction.
Mass The mass of the satellite.
Atmospheric Density Model
  • 1976 Standard. A table look-up model based on the satellite's altitude, with a valid range of 86km - 1000 km.
  • Harris-Priester. Takes into account a 10.7 cm solar flux level and diurnal bulge. Valid range of 0 - 1000 km.
  • Jacchia 1970. The predecessor to the Jacchia 1971 model. Valid range is 90 km - 2500 km.
  • Jacchia 1971. Computes atmospheric density based on the composition of the atmosphere, which depends on the satellite's altitude as well as a divisional and seasonal variation. Valid range is 100km - 2500 km.
  • Jacchia-Roberts. Similar to Jacchia 1971 but uses analytical methods to improve performance.
  • CIRA 1972. Empirical model of atmospheric temperature and densities as recommended by the Committee on Space Research (COSPAR). Similar to the Jacchia 1971 model but uses numeric integration rather than interpolating polynomials for some quantities.
  • MSIS 1986. Empirical density model developed by Hedin based on satellite data. Finds the total density by accounting for the contribution of N2, O, O2, He, Ar and H. 1986 version, valid range of 90-1000 km.
  • MSISE 1990. Empirical density model developed by Hedin based on satellite data. Finds the total density by accounting for the contribution of N2, O, O2, He, Ar and H. 1990 version, valid range of 0-1000 km.
  • NRLMSISE 2000. Empirical density model developed by the US Naval Research Laboratory based on satellite data. Finds the total density by accounting for the contribution of N, N2, O, O2, He, Ar and H. Includes anomalous oxygen. 2000 version, valid range of 0-1000 km. This implementation always calls the gtd7d routine (in contrast to switching between it and gtd7) per the recommendation of Mike Picone, one of the code authors.
  • DTM 2012. The Drag Temperature Model (DTM), 2012 version, is a semi-empirical model which computes the temperature, density, and composition of the thermosphere. Developed at CNES. Valid range of 120 – 1500 km.

  • Jacchia 1970 Lifetime. An implementation of the Jacchia 1970 model that was part of the original Lifetime source code. This was the only density model available for lifetime computations prior to STK 7. Retained for backward compatibility.
Solar Flux File Solar Flux File - An ASCII file containing predicted values of solar flux and geomagnetic activity. The file may follow the format of CSSI long term predicts (.dat), SpaceWeather (.txt) files, or the stkFluxGeoMag (.fxm) files. For more information about the format of these file types, see the Solar Flux Files page of this Help.

The default flux file SolFlx_CSSI.dat used by the Lifetime tool is updated with each new STK release to be the most recently available CSSI file. This file may also be updated using the Data Update utility (DUU).

When the CSSI long term predict file is used by the Lifetime tool, the last solar cycle will be repeated automatically for analyses where the orbit lifetime exceeds the end time of the predicted space weather.

Solar Flux Sigma Level Enter a flux sigma level. The solar flux values used in the orbit lifetime prediction are computed as the nominal solar flux plus the product of the solar flux sigma level and the standard deviation (sigma) associated with the nominal solar flux value. Entering a value of 0 will result in the use of the nominal solar flux values while entering a value of +2 will result in the use of flux levels 2 sigma above the nominal values resulting in shorter lifetime predictions. This setting only affects results when a CSSI predict is specified as the solar flux file.
Advanced... Click this button to define the speed and fidelity of lifetime calculations.
Compute Click this button to perform Lifetime calculations.
SGP4 Compute

The SGP4 Compute button is only available for SGP4 propagated satellites. The SGP4 theory estimates a satellite's orbital lifetime based on the AFSPACECOM SGP4 general perturbations theory. It uses the satellite's 2-line mean elements and thus does not require any of the inputs in the Lifetime nor Lifetime Advanced windows. This theory is a purely analytical solution. It does not provide time-histories of the orbital elements that would be suitable for reports and graphs.

The SGP4 theory uses the BStar value to represent drag effects in 2-line mean element sets. This value cannot be directly converted to a ballistic coefficient (which the standard lifetime algorithm requires). The analytical SGP4 Lifetime algorithm uses the BStar value in the computation of remaining time in orbit. Thus, spurious BStar values can cause large variations in lifetime predictions.

Only use this method if you don't have standard ballistic coefficient information (i.e., when the TLE is the only piece of information given).

Click the SGP4 Compute button to compute the lifetime of the satellite based on this theory.

Report... Generates a report summarizing the satellite's orbital elements over the course of its lifetime. Each element is sampled at perigee passage-the mean, true and eccentric anomalies are always zero and do not display.
Graph... Generates a graph that illustrates the satellite's orbital elements. The graph is especially useful for observing trends and analyzing perturbations to the elements. The changes which the elements undergo are quite complex, especially toward the end of the satellite's life. Generally, though, as a satellite decays you should expect to see the following effects:
  • apogee altitude decreases while perigee altitude remains nearly constant
  • the argument of perigee moves around the orbit plane to the point of minimum atmospheric density
The satellite may be thought of as rotating its apse line and adjusting its eccentricity to extend its life as long as possible.
Show Graphics If ON, the satellite's final orbit displays in the 2D Graphics window. The ground track spans the length of the last orbit and is not intended to represent the exact point of decay.

Options for Defining the Fidelity of Lifetime Calculations

Use the Advanced... button in the Lifetime window to define the speed and fidelity of the calculations to be performed when estimating a satellite's orbital lifetime.

Option Description
Limit Method Specify a method for limiting the run time of the orbit lifetime computation. Options are to limit by a maximum duration of the orbit, by a maximum number of orbit revolutions or by the violation of either limit. Partial results are available for reporting and graphing when the lifetime computation ends due to reaching either limit.
Duration Limit Specify the maximum number of days that will be analyzed before the Lifetime tool stops processing. This limit is used conditionally based on the setting of the Limit Method.
Orbit Count Limit The maximum number of orbits that will be analyzed before the Lifetime tool stops processing. Setting this value to 99999 covers the whole lifetime for most satellites. This limit is used conditionally based on the setting of the Limit Method.
Orbits per Calculation This parameter allows you to directly control the performance of the Lifetime tool. The fewer orbits per calculation, the more precise the lifetime estimate is, but at the expense of compute time. The higher number of orbits per calculation, the less precise the lifetime estimate will be, but calculations are completed much faster. In general, set this parameter to 10 for a quick estimate and 1 for the greatest fidelity.
Gaussian Quadratures Like the previous parameter, this parameter directly affects the performance of the Lifetime tool as well as the accuracy of its results. The drag integration routine is performed by n 9-point Gaussian quadratures per orbit, where n is the number set here. Set this parameter to at least 6 for increased accuracy or lower it for increased speed.
To determine lifetime, the routine approximately integrates the slowly varying orbit elements over time. It does this by integrating over one orbit to determine the rate-of-change of each variable, and then assumes this rate is constant for n number of orbits per calculation (specified on GUI). To increase accuracy, lower this number. The integration of 1 orbit is done over n sub-arcs when n is the number of Gaussian quadratures--the sub-arcs are of constant angular measure (in true anomaly, not time). To increase accuracy, make n larger. Each sub-arc is integrated using a 9-point Gaussian quadrature (i.e., there are 9 sample points within each sub-arc).
Highly eccentric satellites, which are in higher drag regimes for very short times compared with their period, will need to take the number of Gaussian quadratures to be much higher to ensure that drag is sampled adequately.
Decay Altitude The altitude at which the satellite's orbit is determined to be decayed. Lifetime calculations stop at this altitude.
Use 2nd Order Oblateness Correction If ON, a second-order correction is included in the Earth oblateness calculation. This correction is the term, which contributes as much to the gravitational potential as the J3, J4, and J5 terms do.

Second order oblateness effects can result in the modeling of an erroneous growth in the eccentricity for certain orbits, which results in lower than expected lifetime predictions.

Rotating Atmosphere If ON, the west-to-east winds induced by atmospheric rotation are included in the perturbations to the orbit.

Computing Lifetime

Once the appropriate values have been set in the Lifetime and Advanced windows, use the Compute button to start the lifetime calculations. How long the Lifetime tool takes to estimate the satellite's lifetime depends primarily on how high the satellite is at epoch and on the Orbits per Calculation and Gaussian Quadratures parameters.

A Progress window shows the progress of the Lifetime tool and gives you the opportunity to cancel the computations if necessary. The Progress bar reaches 100% when the satellite decays. The "percent of limit" progress message reaches 100% when the number of orbits analyzed by Lifetime equals the Orbit Count Limit parameter set in the Advanced window. If the "percent of limit" significantly outpaces the Progress bar, you may want to cancel the computations and increase the Orbit Count Limit. Otherwise, Lifetime will probably reach the limit before the satellite actually decays. Similarly, if the Progress bar moves very slowly or if the time remaining steadily rises, Lifetime may take a while to estimate the orbital lifetime of the satellite.

For SGP4 satellites, the orbital lifetime can be estimated by either the primary orbit lifetime theory or the SGP4 analytical theory. The SGP4 theory estimates a satellite's orbital lifetime based on AFSPACECOM SGP4 general perturbations theory. It uses the satellite's two-line mean elements and, as such, does not require any of the inputs in either of the Lifetime or Lifetime Advanced windows. As a purely analytical solution, it does not provide time-histories of the orbital elements suitable for reports and graphs. This solution method is also known to be a very poor predictor of re-entry when used during the time period immediately preceding (within a few weeks) the event. Use the SGP4 Compute button to compute the lifetime based on this theory.

If you want a quick estimate, cancel the calculations and adjust the Orbits Per Calculation and/or Gaussian Quadratures fields in the Advanced window. Since the integration of atmospheric drag effects is computationally expensive, reducing the number of Gaussian Quadratures noticeably increases Lifetime's speed. Some accuracy will be lost, but the difference in total lifetime for near-Earth satellites should be small.

Increasing the Orbits Per Calculation parameter can also significantly increase Lifetime's speed. When the number of Orbits per Calculation is greater than one, Lifetime assumes that the perturbations to the satellite's orbit remain constant over the number of orbits specified.

The Lifetime tool runs until either the satellite decays or the Orbit Count Limit is reached. A satellite is assumed to have decayed when its height of perigee drops below 64 km.

Reviewing Lifetime Results

The Lifetime tool estimates the orbital lifetime of a satellite and provides the corresponding date of decay. It should be emphasized that although the Lifetime computations are based on sophisticated orbital theory and accurate environment models, the result is still an estimate. Due to the seemingly random 10% variation in atmospheric density and because of the difficulty in accurately predicting solar activity, satellite lifetimes cannot be determined with accuracy better than +10% of the actual lifetime. Furthermore, assumptions and simplifications made in order to produce a practical computer implementation of the lifetime theory introduce an additional degree of uncertainty in the final result.

The Lifetime tool is not intended to determine an exact time of decay, a specific geographical "impact point," or to what degree a satellite might survive its descent through the atmosphere.