Crew 8 launched to a 191x215km orbit (ref), but the ISS is orbiting at 425km (ref). I'm curious as to the reason why they didn't launch into, for example, a 191x425km orbit initially.

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2 Answers 2


I think this really boils down to a more-general question that hasn't been asked/answered (that I could find before) on this site: why are concentric / coelliptic rendezvous trajectories used instead of direct intercept / direct rendezvous / first apogee?

The short answer is: the coelliptic profile gives you time for on-orbit checkouts, is more tolerant of less-exact maneuvers than a direct intercept, and gives you easier control over time of docking. It also gives crew a chance to rest after a long launch day before rendezvous activities.

To put that into the context of this question, see the following image that I've excerpted from "SpaceX Crew Dragon's flight plan for NASA Astronauts’ debut voyage" which is itself from a NASA broadcast (if anyone could help me get a primary source--I assume it showed up in the Demo 2 broadcast--I'd appreciate it). This shows the evolution of the trajectory; the initial orbit you describe is the lower orbit that intersects Mark 02, though the labeling and is maybe not what I'd have done myself. screen capture from a NASA broadcast showing the various maneuvers and intermediate orbits from launch to ISS rendezvous for Crew Dragon

Compare that to the following excerpts from "Gemini VI Profile Selection" from JSC 63400 History of Space Shuttle Rendezvous, Rev 3, page 20:

Figures 3.1, 3.2, and 3.3 from History of Space Shuttle Rendezvous, showing tangential, coelliptic, and direct-intercept rendezvous respectively. Tangential is several elliptical orbits evolving into the desired target elliptical orbit. Coelliptic uses an intermediate orbit to match inclination, catch up with, and then intercept. Direct intercept is the OP's proposed match-after-half-an-orbit ascent.

Tangential Orbit (Mission Plan 1) - This profile involved launching the Gemini spacecraft into an elliptical orbit tangential to the Agena Target Vehicle (ATV) orbit (Figure 3.1). Rendezvous would occur near the apogee of the fourth Gemini orbit. However, this technique did not guarantee proper lighting conditions or consistent relative dynamics in the terminal phase under dispersed conditions.

Concentric Flight Profile (Mission Plan 2) - This used the same maneuver plan as the ground targeted phase in the tangential orbit profile, but had a different terminal phase. Rather than ground targeting placing the spacecraft on an intercept trajectory, it placed the Gemini in a co-elliptic orbit with respect to the target spacecraft (Figure 3.2). The intercept maneuver, Terminal Phase Initiation (TPI), was executed while the chaser vehicle was on an orbit coelliptic with the Agena. The length of the co-elliptic phase could be controlled to ensure appropriate lighting during the terminal phase and adequate coverage by ground tracking. The terminal phase would begin once a trajectory criterion was met.

First Apogee or Direct Rendezvous (Mission Plan 3) - The Titan II booster would place the Gemini spacecraft on an intercept trajectory with the Agena (Figure 3.3). Gemini would achieve radar lock on the target soon after orbit insertion. However, the short amount of time for the crew to conduct on-orbit checkout of Gemini systems and rendezvous procedures made the timeline impractical. Furthermore, the trajectory was highly sensitive to ascent dispersions and liftoff delays. Trajectory dispersions would have to be corrected by the on-board system, without help from ground tracking. In case of a dispersed trajectory that made rendezvous impossible, a backup rendezvous profile was needed.

After an extensive trade study, a coelliptic rendezvous profile was chosen for execution on Gemini VI at a meeting on June 15, 1964 (Figure 3.2). The length of the coelliptic phase permitted control over terminal phase lighting, and provided a terminal phase that was less sensitive to trajectory dispersions than the direct rendezvous and tangential orbit profiles.

Of the three concentric provided the most flexibility, had a terminal phase that was the least sensitive to dispersions and facilitated easier definition of backup procedures. This helped ensure standardized crew procedures and training, even with mission-to-mission variations in the pre-terminal phase rendezvous profile. Furthermore, compared to the direct rendezvous profile, the crew did not have to conduct rendezvous activities during the first orbit, as it was preferred to spend the first orbit conducting spacecraft systems checks. The coelliptic approach also facilitated the use of manual backup guidance techniques in the event of system failures (sensor failure, computer failure, or loss of communications with Mission Control).

Emphasis mine on the specific difficulties of the direct ascent to the target orbit. I may be over-generalizing to say that variations of the concentric / coelliptic rendezvous trajectory have been flown ever since for the reasons described in the excerpt; further such trajectories are evident in e.g. Figure 9.1 from the same source. In my career, though I was only directly concerned with the nearest-field rendezvous details, I saw many such rendezvous trajectories for various vehicles. I was also in a single meeting where I believe they were mis-applied out of habit in trajectory design for rendezvous with the Lunar Gateway in its NRHO (I specifically remember thinking "wow, that's a really flat orbit").

Other rendezvous trajectories are available, and there was speculation on this StackExchange that Dragon 2 would employ the expedited 6-hour rendezvous as demonstrated by Progress and Soyuz, but to my knowledge that has never occurred. I'm not aware if there's any technical reason not to or if the NASA customers just continually demanded that things be done the way that they're used to. I'm also not 100% on the details of that trajectory; it may resemble the coelliptic more closely than I think. Nonetheless, I think this answer applies.

  • $\begingroup$ Soyuz flew its first crewed fast track mission in 2013, expediting rendezvous from the previous two days to six hours as you mentioned. What is interesting to note is that Soyuz flew space station missions for 42 years before making this change. Presumably part of what enabled it was advances in spacecraft navigation and propulsion systems. It would be another seven years before Soyuz changed to their current three-hour rendezvous method in 2020. $\endgroup$ Commented Mar 11 at 2:08
  • $\begingroup$ I remember when fast track started cosmonauts mentioned that it makes for a very long day. For any mission, launch day starts many hours before launch for the crew. With the two-day method their rest period began not too long after reaching orbit (I think maybe a couple of hours). With the new method they start their day at the same time, but they now remain active for the six (now three) hour rendezvous period, followed by docking, welcome, safety tour, group video broadcast, etc. And I'm guessing that similar to Shuttle they stay up well into their sleep period on their first day on Station. $\endgroup$ Commented Mar 11 at 2:12
  • $\begingroup$ @StevePemberton excellent point, and one that I think should be easy to get a primary source for, definitely for Crew Dragon, maybe even for Crew 8. I'll look for one to edit into the answer $\endgroup$
    – Erin Anne
    Commented Mar 11 at 2:40
  • $\begingroup$ Interesting. Now that I understand the terminology better, my question could be better stated as "What is the advantage of Concentric over Tangential?" I did not understand the explanation given by the referenced source: "Tangential ... did not guarantee proper lighting conditions or consistent relative dynamics in the terminal phase under dispersed conditions." I now understand the issues associated with First Apogee though. +1 $\endgroup$
    – phil1008
    Commented Mar 11 at 6:10
  • $\begingroup$ I think I have deduced why tangential "...did not guarantee proper lighting conditions ... under dispersed conditions". Still not sure what they meant by "consistent relative dynamics" though. $\endgroup$
    – phil1008
    Commented Mar 11 at 6:30

During SpaceX's CRS-30 broadcast, at T+00:15:30, mission control in Houston provided some information that helps to explain the typical mission profile.

The activation and rendezvous phase of the mission begins after separation from the Falcon 9 and ends with the completion of the final coelliptic burn.

The initial orbit (for the CRS-30 mission) was 190 km by 210 km.

Over the next day and a half, Dragon initiated a series of five major burns to gradually raise its orbit to align more closely with the station.

enter image description here

The faint numbers on the lefthand side of the figure show how much higher the Space Station's orbit is than Dragon's orbit.

A series of burns beginning with the "phase burn" (which is initiated at apogee to raise the orbit's perigee) gradually bring Dragon's orbit closer to the Stace Station's orbit. While the burns all have unique names, they are collectively referred to as "phasing burns".

These burns are all done with the forward bulkhead Draco engines which are located on the tip (or nose) of the Dragon spacecraft.

SpaceX also describes the set of phasing burns as a

"methodical approach" to the space station

implying that they are designed to prioritize a high probability of mission success over arriving at the station quickly.

  • $\begingroup$ Exactly. We have reliable GPS now, SpaceX has a best-part-is-no-part philosophy, and, IIRC, the Soyuz missions go to the ISS more directly. Wouldn't the best phasing burns be no phasing burns? Is it not worth doing or is the system we're using not rewarding innovations? $\endgroup$
    – phil1008
    Commented Mar 23 at 1:00
  • 1
    $\begingroup$ The timing on a direct-intercept launch is really tight, and launch windows don't come up very often. $\endgroup$
    – Mark
    Commented Mar 23 at 3:21
  • $\begingroup$ To put some hard numbers on it, if you want to launch directly to "approach initiation", there's roughly one launch window every eight years. If you're going for "final coelliptic burn", it's three windows every two years. $\endgroup$
    – Mark
    Commented Jun 14 at 21:52

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