A launch vehicle failure stranded ATS-4 in a much lower than planned orbit, making the satellite nearly useless. Although the spacecraft's systems were functional, little data was obtained. Spacecraft: Gravity gradient stabilized, 130 W BOL. Payload: C-band communications package

Design Life: 3 years. Typical orbit: Agena failure left it in transfer orbit. Length: 1.80 m (5.90 ft). Maximum Diameter: 1.40 m (4.50 ft). Mass: 391 kg (862 lb).

• 1968 August 10 - ATS 4 - Program: ATS. Launch Site: Cape Canaveral. Launch Complex: LC36A. Launch Vehicle: Atlas. FAILURE: Centaur oxidizer leak. No restart. Mass: 391 kg (862 lb). Perigee: 220 km (130 mi). Apogee: 769 km (477 mi). Inclination: 29.10 deg. Period: 94.50 min.
  Applications Technology Satellite that was to have been put into a geosynchronous transfer orbit, instead was left in a nearly-useless LEO orbit. ATS-4 included two cesium contact ion engines. Flight test objectives were to measure thrust and to examine electromagnetic compatibility with other spacecraft subsystems. The 5 cm diameter thrusters were designed to operate at 0.02 kW and provide about 89 microN thrust at about 6700 s specific impulse. The thrusters had the capability to operate at 5 setpoints from 18 to 89 microN. Thrusters were configured so they could be used for East-West station-keeping. Prior to launch, a 5 cm cesium thruster was life tested for 2245 hours at the 67 microN thrust level. However the Centaur upper stage did not achieve a second burn and the spacecraft remained attached to the Centaur in a 218 km by 760 km orbit. It was estimated that the pressure at these altitudes was between 10^-6 and 10^-8 Torr. Each of the two engines was tested on at least two occasions each over the throttling range. Combined test time of the two engines was about 10 hours over a 55 day period. The spacecraft re-entered the atmosphere on October 17, 1968. TheATS-4 flight was the first successful orbital test of an ion engine. There was no evidence of IPS electromagnetic interference related to spacecraft subsystems. Measured values of neutralizer emission current were much less than the ion beam current, implying inadequate neutralization. The spacecraft potential was about -132V which was much different than the anticipated value of about -40V.