The　communication　and　observation　spacecraft　generally　consists　of　the　spacecraft　body,　the　composite　connecting　beams,　the　working　payload　and　the　CMG/momentum　wheel　for　integrated　attitude-vibration　control.　When　such　complex　space　structure　operates　in　near-Earth　orbit,　the　sudden　change　in　thermal　load　may　cause　a　solar　flux　shock　to　the　structure,　which　is　likely　to　excite　the　flutter　motion　of　the　structure.　The　flutter　mechanism　of　a　rigid-flexible　coupled　composite　space　structure　configured　with　the　momentum　wheel　is　investigated.　In　this　study,　the　heat　conduction　equation　of　the　composite　beam　appendage　is　derived.　Then,　the　dynamic　equations　of　the　system　influenced　by　the　momentum　wheel　and　composite　material　parameters　are　obtained　based　on　the　Kane＇s　equations.　Using　the　Routh-Hurwitz　stability　criterion,　the　stability　boundaries　of　the　flexible　appendage　are　obtained　for　different　composite　material　parameters,　different　operating　conditions　of　the　momentum　wheel,　and　different　attitude　angular　velocities.　Finally,　the　flutter　behavior　is　analyzed　based　on　the　dynamic　response　of　the　flexible　structure　and　the　flutter　stability　boundaries　are　verified.