This study presents a new ultra-high temperature composite fabricated by using zirconium diboride (ZrB₂), silicon carbide (SiC), and tantalum carbide (TaC) with the volume ratios of 70%, 20%, and 10%, respectively. To attain this novel composite, an advanced processing technique of spark plasma sintering (SPS) was applied to produce ZrB₂–SiC–TaC. The SPS manufacturing process was achieved under pressure of 30 MPa, at 2000 °C for 5 min. The micro/nanostructure and mechanical characteristics of the composite were clarified using X-ray diffraction (XRD), field-emission scanning electron microscopy (FESEM), and nano-indentation. For further investigations of the product and its characteristics, X-ray fluorescence (XRF) analysis and X-ray photoelectron spectroscopy (XPS) were undertaken, and the main constituting components were provided. The composite was densified to obtain a fully-dense ternary; the oxide pollutions were wiped out. The mean values of 23,356; 403.5 GPa; and 3100 °C were obtained for the rigidity, elastic modulus, and thermal resistance of the ZrB₂–SiC–TaC interface, respectively. To explore the practical application of the composite, the natural frequency of an aircraft wing considering three cases of materials: (i) with a leading edge made of ZrB₂–SiC–TaC; (ii) the whole wing made of ZrB₂–SiC–TaC; and (iii) the whole wing made of aluminum 2024-T3 were investigated employing a numerical finite element model (FEM) tool ABAQUS and compared with that of a wing of traditional materials. The precision of the method was verified by performing static analysis to obtain the responses of the wing including total deformation, equivalent stress, and strain. A comparison study of the results of this study and published literature clarified the validity of the FEM analysis of the current research. The composite produced in this study significantly can improve the vibrational responses and structural behavior of the aircraft’s wings.