Orbital Pegasus - Manual (Page 25)

Release 7.0 Apr 2010 14
Pegasus User’s Guide
loads. It shall survive those conditions in a
manner that ensures safety and that does not
reduce the mission success probability. The
primary support structure of the spacecraft shall
be electrically conductive to establish a single
point electrical ground. Spacecraft design loads
are defined as follows:
 Design Limit Load — The maximum predicted
ground-based, captive carry, or powered flight
load, including all uncertainties.
 Design Yield Load — The Design Limit Load
multiplied by the required Yield Factor of
Safety (YFS) indicated in Figure 4-1. The
payload structure must have sufficient strength
to withstand simultaneously the yield loads,
applied temperature, and other accompanying
environmental phenomena for each design
condition without experiencing detrimental
yielding or permanent deformation.
 Design Ultimate Load — The Design Limit
Load multiplied by the required Ultimate
Factor of Safety (UFS) indicated in Figure 4-1.
The payload structure must have sufficient
strength to withstand simultaneously the
ultimate loads, applied temperature, and other
accompanying environmental phenomena
without experiencing any fracture or other
failure mode of the structure.
4.2. Payload Testing and Analysis
Sufficient payload testing and/or analysis must be
performed to ensure the safety of ground and
aircraft crews and to ensure mission success. The
payload design must comply with the testing and
design factors of safety in Figure 4-1 and the FAA
regulations for the carrier aircraft listed in the
CFR14 document, FAR Part 25. UFS shown in
Figure 4-1 must be maintained per Orbital SSD
TD-0005. At a minimum, the following tests must
be performed:
 Structural Integrity — Static loads or other
tests shall be performed that combine to
encompass the acceleration load environment
presented in Section 4.3. Test level
requirements are defined in Figure 4-1.
 Random Vibration — Test level requirements
are defined in Figure 4-2.
4.3. Payload Acceleration Environment
Maximum expected loads during captive carry and
launch are shown in Figures 4-3, 4-4, and 4-5.
The Pegasus air-launch operation results in a
launch vehicle/OCA separation transient at drop.
The drop transient acceleration limits presented
here are based on two assumptions:
(1) Pegasus Standard 23” or 38” payload
separation system is used.
(2) The first fundamental lateral frequency of the
spacecraft cantilevered at the payload
interface (excluding the payload separation
system) is greater than 20 Hz.
If either assumption is violated, mission-specific
analyses are required. For all missions, accurate
estimation of the drop transient loading requires a
coupled loads analysis (CLA), which uses Orbital
and customer-provided finite element models to
predict the transient environment (see Section
8.3.3 for details).
Transient loading also exists due to motor ignition.
Stage 1 provides the worst-case loading due to
motor ignition. The Stage 1 ignition acceleration
limits at the payload interface are listed in
Figure
4-3. The Stage 1 shock response
spectrum (SRS) at the payload interface is shown
in Figure 4-6. As is the case with the drop
transient, accurate estimation of loading requires a
CLA. The Stage 1 ignition transient CLA requires
finite element models of the Pegasus avionics
structure, payload separation system, and the
payload.
4.4. Payload Random Vibration Environment
The maximum expected random vibration levels at
the payload interface are shown in Figure 4-7.
Random vibration data recorded during multiple
Pegasus missions was used to create this overall