Summary of Failure Analysis of SSLV-D1 Mission & Recommendations for SSLV-D2

Small Satellite Launch Vehicle D1

The SSLV-D1 was the first developmental mission of this new launch vehicle. The objective of development missions is to prove the launch vehicle design and architecture and to bring out any residual unknowns not identified in the qualifications tests and analysis during its development journey. SSLV-D1 mission demonstrated the satisfactory integrated performance of SSLV in all its systems including its flight through the aerodynamic regime, which is an accomplishment by itself.

Cause of Anomaly

Subsequent detailed analysis of the flight events and observations ranging from countdown, lift-off, propulsion performance, stage separations and satellite injection revealed that there was a vibration disturbance for a short duration on the Equipment Bay (EB) deck during the second stage (SS2) separation, that affected the Inertial Navigation System (INS), resulting in declaring the sensors faulty by the logic in Fault Detection & Isolation (FDI) software

SSLV Inertial Navigation System

SSLV uses a newly developed inertial navigation system, MEMS Inertial Navigation System or MINS-6S, which consists of 6 MEMS Gyros (for measuring rotation rates) and 6 Ceramic Servo Accelerometers (for measuring accelerations) assembled in a vibration isolated temperature-controlled assembly. The system is also designed with in-built NavIC receiver and also acts as the MINS navigation computer processor for running Inertial Navigation and Aided Navigation software. A novel algorithm estimates the error in the attitude (orientation) introduced due to the MEMS Gyros, position & velocity of the MEMS INS, using the NavIC data and corrects them, so that mission accuracy is achieved. Knowing the health of accelerometer is of paramount importance for the functioning of MINS, as it is used for attitude aiding. The failure detection logic identifies degraded accelerometer (one or all of 6) and isolate the same for improved mission performance.

During the SS2 separation event, all the six accelerometers inside the MINS package experienced measurement saturation due to high vibration levels for a short duration. The accelerometers got saturated at different time instants, within the 20 millisecond (ms) interval of data sampling, which resulted in different acceleration values being measured by each sensor. This resulted in high residue values (difference among them) beyond a specified limit for a duration of 2 seconds. The software implementing the Fault Detection & Isolation (FDI) assessed that the sensor outputs crossed the pre-set threshold limit and raised the alert/flag for the salvage mission mode initiation, which is a safety approach for mission. However, the accelerometers were found functioning well after this transient event. The salvage mission mode got fully executed without the support of the accelerometer data and injected the satellite to an unstable orbit due to lower injection velocity (~56m/s less than the required 7693 m/s). Though this is as programmed and expected, if kept observing for longer duration, the residue among the sensors would be within limit and failure logic would not have been executed.

Shock And Vibrations During Stage Separation

The vibrations at the satellite interface as measured during the flight, were well within the pre-flight test specifications except during second stage (SS2) separation event when the Equipment Bay and Satellite interface had dominant vibration response in the first axial mode. The shock response measured at EB during the SS2 separation exceeded the expectations and ground tests levels both in low and high frequency as well as in time duration.

Due to this shock, excitation at the MINS sensors was persistent for more than 10 milli second (ms) duration which was not expected. It is generally observed that shock from such events last for about 2 ms, whereas here a shock of 2-3 ms duration and subsequent oscillations lasting for more than 10 ms was observed.

Further, based on the flight telemetry data, all the accelerometers were found functioning normally after the transient event till the end of the mission, indicating no damage to them. However due to mission salvage setting by the FDI program, data from these accelerometers were not used for further mission execution.

Satellite Injection Into An Unintended Orbit

Though the salvage mode was initiated with the purpose of saving the mission, it could not inject the satellites to a safe orbit. The third stage, SS3 ignition was commanded by the sequence program. Subsequently, the vehicle was guided through time-based open loop mode steering without feedback, as the accelerometer data was declared faulty. At the end of SS3 burnout, the satellites were separated safely as programmed. In the implemented salvage scheme, with predefined time-based open loop guidance scheme, the vehicle attitude reference could be erroneous depending on the MEMS Gyro errors. Further, there is no knowledge of the actual velocity of the vehicle as well, since the velocity is computed from accelerometer data. VTM ignition was bypassed as programmed, since it could be a deterrent to the success of salvage option in some cases. The shortage of about 56 m/s at the end of SS3 burn out in final velocity (due to cumulative deficiency in performance of all propulsion stages) and loss in pointing accuracy due to sensor errors, the targeted orbit could not be achieved. This indicates that execution of salvage option in all situations need not always lead to successful placement of satellites in an orbit.

Recommendations & Correction Actions

Change of Separation System

The separation system used for separating Second stage from third stage was based on a Circular Expanding Bellow system which shear the rivets and give axial separation velocity. This system is replaced by well proven Marman band system for separation and springs for giving axial separation velocity. The new system is proven to be generating lesser shock and is already used in the separation of third stage.


The FDI logic based on the accelerometer threshold is modified to evolve a more realistic approach based on the data generated through system level tests, integrated separation tests and flight. The accelerometer residue logic checking in MINS is modified to handle transient events. Moving average window is modified so that in case of failure identification of multiple sensors in MINS, checking for a longer duration before setting the salvage mode is implemented.

Dynamic Characterization And Design Modification of Structures

Dynamic assessment of EB & Satellite assembly along with VTM is carried out and structural design modified to increase the frequency of the structures. Modifications in EB deck & Satellite deck were implemented to minimize response to the observed excitations.

Usage of NaVIC Data

Further, in case of failure of inertial system sensors, the mission will be progressing using NavIC data in a closed loop guidance scheme.

VTM Will Be In Loop For Salvage Mode

In case of failure of inertial sensors and non-availability of NavIC data (for more than 10 sec), an open loop steering guidance will be executed. The propulsive capability of VTM will be considered in this salvage mode also and thrusters will be operated to ensure the minimum required perigee for the mission.