Mission Radar Radome Installation - Structural Substantiation

INTRODUCTION

Structural substantiation of the DHC-8-100 mission-radar radome installation, tracing aerodynamic loads through the perimeter attachments into a 16-ply fiberglass flange and the aircraft lower fuselage. The work combines conservative aerodynamic loading, composite-joint allowables, 3D rigid-body fastener distribution, Classical Laminate Plate Theory (CLPT), and local interface shear checks.

AIRCRAFTDHC-8-100Mid-fuselage external radome

PRIMARY INTERFACE16-ply fiberglass flangeMeasured cured thickness: 0.192 in

LOAD PATHAerodynamics → Attachments → FlangeEight-bolt bounding reaction model

SUBSTANTIATIONJoints + CLPT + ShearPositive margins / inspection checks

Aircraft station diagram with radome installation location and dimensions — Aircraft-station definition and radome installation envelope.

DESIGN ASSESSMENT

RADOMESupplier-procured aerodynamic fairing

Protects mission radar/communications equipment while maintaining the required external contour and low signal attenuation.

CUSTOM FLANGEPrimary structural interface

In-house fiberglass laminate bridges the supplier radome to the aircraft lower skin and spreads attachment load around the perimeter.

FITTING SUB-ASSEMBLYMechanical load transfer

Transfers local attachment reactions between the radome installation and surrounding fuselage structure.

RADOME FRAMESPort + starboard support

Receive the remaining installation loads through the fitting architecture.

Radome installation schematic and primary structural interfaces.

LOAD-PATH RATIONALE

Separate global aerodynamic demand from local composite-interface checks.
External lift and drag enter through the radome shell, resolve into perimeter attachment reactions, pass through the custom fiberglass flange as in-plane bearing and out-of-plane pull-through demand, and then transfer into the aircraft lower-skin/frame structure. Treating the global aerodynamics, discrete fastener reactions and local laminate response as separate calculation stages keeps the load path traceable while avoiding unnecessary detail in any one model.

Fiberglass flange layup and radome interface drawing — Fiberglass flange layup and radome interface detail.

CURED THICKNESS0.192 in

REINFORCEMENTFiberglass Style 7781MIL-C-9084

MATRIXEPOCAST 50A + 9816Manufacturer mixing/cure schedule

LAYUP16 plies0/45/90/90/45/0/0/45/90/90/45/0/0/90/90/0°

MATERIAL-SELECTION RATIONALE

Fiberglass supports the electromagnetic function of the radome

Fiberglass provides the required structural interface while avoiding the radar-attenuation concerns associated with a metallic integration flange.

MATERIAL PROPERTIES

ENVIRONMENTAL BASIS

Hot-wet strength set

Unknown fabrication condition is bounded with the conservative 160°F / 95% RH laminate strength set.

UNCERTAINTY ALLOWANCE

Additional 25% knockdown

Source allowables are reduced a further 25% to cover manufacturing and final-laminate uncertainty.

F_tu132.11 ksi

F_tu225.75 ksi

F_cu136.14 ksi

F_cu236.14 ksi

F_su128.44 ksi

Strength basis: [Analysis & Design of Composite & Metallic Flight Vehicle Structures by Richard Abbott], aligned by the source with the lower bound of MIL-HDBK-17-2F.

ILSS BASIS

6.5 ksi typical → 4.875 ksi design value

The source treats glass/epoxy interlaminar shear as comparable to a carbon laminate using the same resin and then applies the same 25% uncertainty reduction to the 6.5 ksi typical value.

WHY THIS IS CONSERVATIVE

Lower-bound manufacturing sensitivity retained

ILSS is sensitive to cure quality and laminate processing. Using the knocked-down value deliberately avoids crediting a best-case fabrication state that is not documented.

JOINTS ALLOWABLE

Joint capacity is governed by the weakest applicable failure mode for the actual fastener/laminate geometry: fastener shear versus laminate bearing in-plane, and fastener tension versus laminate pull-through out-of-plane.

Joint allowable philosophy

Separate the in-plane and out-of-plane load paths, then govern each with the minimum applicable capacity.

P_{in-plane}=\min(P_{fastener\ shear},P_{laminate\ bearing})

P_{out-of-plane}=\min(P_{fastener\ tension},P_{pull-through})

COMPOSITE-JOINT METHODOLOGY Stay inside the demonstrated test envelope.

Check t/D, e/D, fastener diameter, and reference-laminate thickness against the demonstrated test envelope, then apply only the required knockdowns before calculating bearing capacity.

Joint’s Bearing Strength:

GLASS-FIBER BASELINEF_bru = 40 ksiWorst-environment general value from the source methodology.

Geometry window0.5 ≤ t/D ≤ 2e/D > 3.0

If outside the windowUse source-chart knockdownsScale for smaller fastener diameter / thinner laminate when applicable

MODELING RATIONALEDo not credit composite bearing strength beyond the test envelope.

Chart-based knockdowns are used whenever geometry departs from the published test configuration, preventing unsupported extrapolation of the 40 ksi baseline bearing strength. [Analysis & Design of Composite & Metallic Flight Vehicle Structures by Richard Abbott – Section 12.2.4.2].

Joint’s Pull-Through Strength:

Pull-through is evaluated with the source semi-empirical method correlated to interlaminar shear strength; a glass-fiber reinforcement factor of 0.75 is used. [Analysis & Design of Composite & Metallic Flight Vehicle Structures by Richard Abbott].

F_{pull} = \frac{2\alpha t\pi D_{f}\tau_{\max}}{2.9 - 0.018\frac{t}{D_{s}} - 0.51\left( \frac{t}{D_{s}} \right)^{2}}

t

Laminate thickness

D_s

Bolt shank diameter

D_f

Bolt head / washer diameter

\tau_{max}

Interlaminar shear stress

α

Reinforcement factor: 1.0 carbon, 0.75 glass

Fastener Joint Capacity

AN3-13A and AN525-10R16 fastener properties [Technical Data Sheet-NASM3-20, NASM525].

P/N	Type	Head	Size (Callout)(Thread)(Length)	Material	D_s (in)	(F_su)(F_tu) for fastener material [ksi]	(f_su)(f_tu) for fastener in single shear [lb_f]	Thread Standard
AN3-13A	Aircraft Bolt	Hex	(#10‑32)(UNF‑3A)(1.40625)	Non-Corrosion Resistant Steel	0.19	N/R	(2125)(2210)	MIL-S-7742
AN525-10R16	Screw	Recessed	(#10-32)(UNF-3A)(1)	Cadmium Plated Alloy Steel	0.19	(0.6x125)(125)		MIL-S-7742

AN3-13A

Aircraft Bolt

Hex

(#10‑32)(UNF‑3A)(1.40625)

Non-Corrosion Resistant Steel

0.19

N/R

(2125)(2210)

MIL-S-7742

AN525-10R16

Screw

Recessed

(#10-32)(UNF-3A)(1)

Cadmium Plated Alloy Steel

0.19

(0.6x125)(125)

MIL-S-7742

AN3-13AGoverning in-plane: 1,109 lb_fComposite bearing governs over 2,125 lb_f fastener shear.Governing out-of-plane: 1,634.65 lb_fPull-through governs over 2,210 lb_f fastener tension.

AN525-10R16Governing in-plane: 1,109 lb_fSame laminate bearing geometry governs.Governing out-of-plane: 1,634.65 lb_fSame washer / laminate pull-through geometry governs.

Representative governing joint calculation

t/D = 0.192/0.19 = 1.01 → no thickness knockdown.
Using a 50% reduced edge distance gives e/D = 0.675/0.19 = 3.55 → no edge-distance knockdown.
The 0.19 in shank versus the 0.25 in reference diameter requires a 0.76 factor, reducing Fbru to 30.40 ksi and giving a laminate bearing capacity of 1,109 lbf.

Bearing governs the in-plane joint allowable at 1,109 lbf after the 0.76 diameter knockdown (Fbru = 30.40 ksi).

With a 0.875 in AN970-3 washer and the conservative 4.875 ksi ILSS design value, the pull-through calculation gives: [Technical Data Sheet-AN970-3 Washer].

f_{pull}=\frac{2\alpha t\pi D_f\tau_{max}}{2.9-0.018\frac{t}{D_s}-0.51\left(\frac{t}{D_s}\right)^2}

f_{pull}=\frac{2\pi(0.75)(0.192)(0.875)(4.875\times10^3)}{2.9-0.018\left(\frac{0.192}{0.19}\right)-0.51\left(\frac{0.192}{0.19}\right)^2}

f_{pull}=1{,}634.65\;lb_f

Pull-through is lower than fastener tension, so 1,634.65 lbf governs the out-of-plane joint allowable.

Governing ultimate joint capacities.

Joint	In Plane Strength [lb_f]	Out of Plane Strength [lb_f]
AN3-13A Fasteners’ Joints	1,109	1,634.65
AN525-10R16 Fasteners’ Joints	1,109	1,634.65

AN3-13A Fasteners’ Joints

1,109

1,634.65

AN525-10R16 Fasteners’ Joints

1,109

1,634.65

LOADS

IDEALIZATIONInverted 2D wingRadome fairing treated as a downward-lift surface.

SEA-LEVEL DENSITYρ = 0.00238 slug/ft³

MAX OPERATING SPEEDV_MO,max = 242 kt408.45 ft/s

AERODYNAMIC CENTER1/3 chordMid-height in z

LOAD-CASE SELECTION RATIONALE

Use sea-level density and maximum operating speed, bound radome lift with the aircraft-wing cruise coefficient, use the cited maximum drag coefficient, and apply a 1.5 ultimate factor. Level fuselage attitude bounds the downward-lift case for the inverted fairing.

Radome fairing dimensions and assumed aerodynamic-center location.

Aerodynamic Center (in)	Dimensions (in)
X_c.g.	Y_c.g.	Z_c.g.	L	W	D
41.33	0	-13.985	124.33	65.72	27.97

X_c.g.

Y_c.g.

Z_c.g.

41.33

-13.985

124.33

65.72

27.97

Radome fairing dimensions and aerodynamic center — Radome fairing dimensions and aerodynamic-center definition.

LIFT BOUND

Horizontal fuselage attitude governs the inverted fairing

Level fuselage attitude bounds downward lift; positive angle of attack reduces the effective lift of the inverted fairing.

DRAG BASIS

C_D,max = 0.051

Maximum radome drag coefficient is taken from the cited Horner drag data [Fluid-Dynamic Drag-Horner, page 8-5, Fig. 11].

Conservative lift-coefficient bound:

C_{L,wing}=\frac{34,500}{\frac12(0.00238)(408.45)^2(604)}=0.2877

C_{L,rad,max}=C_{L,wing}=0.2877

Ultimate radome lift:

L_{max}=\frac12\rho V^2SC_L=1,944.66\ lb_f

L_{ult}=1.5(1,944.66)=2,916.99\ lb_f

The same dynamic-pressure basis is applied to drag using $C_{D,max}=0.051$ :

D_{max}=\frac12\rho V^2SC_D=129.248\ lb_f

D_{ult}=1.5(129.248)=193.87\ lb_f

ULTIMATE AERODYNAMIC LOADSBounding global demand

Downward lift 2,916.99 lb_f

Drag 193.87 lb_f

Factor 1.5

STATIC ANALYSIS

Fasteners' Reaction Forces

BOUNDING ATTACHMENT MODELCredit only eight extreme attachment bolts.

Aerodynamic force and moment are reacted at selected FWD, starboard, AFT and port extreme locations using 3D rigid-body analysis. Concentrating the global load into this sparse attachment set is intentionally conservative relative to distributing it around the full perimeter.

Radome eight-bolt attachment reaction model — Eight extreme attachment locations used for the bounding 3D rigid-body reaction model.

Shear and axial reactions [lb_f] at the assumed main attachment points.

Fastener No	Coordinates	f_x	f_z	Shear	Axial
1	0, 0, 0	-34.2	461.2	34.2	461.2
2	17.47, 16.43, 5	-24.2	426.5	24.2	426.5
3	34.94, 32.86, 10	-14.3	391.7	14.3	391.7
4	79.63, 16.43, 5	-24.2	302.8	24.2	302.8
5	124.33, 0, 0	-34.2	213.8	34.2	213.8
6	79.63, -16.43, 5	-24.2	302.8	24.2	302.8
7	34.94, -32.86, 10	-14.3	391.7	14.3	391.7
8	17.47, -16.43, 5	-24.2	426.5	24.2	426.5

GOVERNING ATTACHMENT DEMANDExtreme-bolt reaction

Max shear 34.2 lb_f

Max axial 461.2 lb_f

PASS

{M.S}_{s} = \frac{1,109}{34.2 \times 1.15} - 1

{M.S}_{t} = \frac{1,634.65}{461.2 \times 1.15} - 1

\boxed{{M.S}_{s} = 27.2\ \ }

→ PASS

\boxed{{M.S}_{t} = 2.08\ \ }

→ PASS

Flange

METHOD

Classical Laminate Plate Theory

The 16-ply flange is represented through the ABD stiffness matrix, with ply-level direct, bending and total stresses recovered through the laminate.

FAILURE CRITERIA

Tsai-Wu · Hill · Hoffman

The source spreadsheet compares ply stresses against multiple composite failure criteria, allowing the minimum margin to identify the governing interaction.

Governing leading-edge attachment input to the flange CLPT check.

Location No	f_x	Moment (M_y) due to f_z
1	34.2	461.2x2.70=1,245.24

34.2

461.2x2.70=1,245.24

Critical flange plate element — Critical 3.7 × 2.7 in flange plate element at the governing leading-edge attachment.

CRITICAL-ELEMENT RATIONALEAnalyze the flange where local attachment demand is highest.

A 3.7 × 2.7 in leading-edge plate element is isolated at the highest attachment demand; drag and the attachment moment are converted to running loads for CLPT input.

Laminate properties and ABD matrix — Laminate input, material properties and ABD stiffness matrix used in the CLPT analysis.

Ply stresses and laminate margins — Ply direct, bending and total stresses with the minimum laminate margins.

FLANGE GOVERNING RESULTHill criterion

Minimum margin 13%

Material knockdown 25%

PASS

The reported positive margin is obtained after the 25% laminate-property knockdown used to cover source manufacturing uncertainty.

Local radome-to-flange interface shear:

Local radome-to-flange shear element — Local 2.5 × 2.5 in radome-to-flange element used for the interface shear check.

{f_{s,x} = \frac{D}{{Height}_{Radome\ }} \times Z_{Element\ } }{= \frac{193.87}{27.97} \times 2.5 = 17.33\ {lb}_{f}}

{f_{s,z} = \frac{L/2}{{Length}_{Radome\ }} \times X_{Element} }{= \frac{2,916.99/2}{124.33} \times 2.5 = 29.33\ {lb}_{f}}

f_{s} = \sqrt{{17.33\ }^{2} + {29.33}^{2}} = 34.06\ {lb}_{f}

Average shear stress over the 2.5 × 2.5 in interface element:

F_s=\frac{34.06}{2.5\times2.5}=5.45\ psi

Interface shear - PASS5.45 psi demand versus 270 psi minimum EPOCAST 50-A1 lap-shear strength [Epocast 50-A1 Resin - Hardener 9816].

All other structural components

INSPECTION CHECKRemaining fitting and frame structure

The source closes the remaining components by inspection because the attachment tension and shear reactions are low; consequently, the loads passed through the tension-fitting sub-assembly into the radome frames are also small.

PASS BY INSPECTION

REFERENCES

Composite Materials & Analysis Methods

Composite Materials Handbook - MIL-HDBK-17 / CMH-17
Analysis & Design of Composite & Metallic Flight Vehicle Structures - Richard Abbott
The Strength of Bolted Joints in Multidirectional CFRP Laminates

Joint & Fastener Data

Fastener Design Manual - NASA Reference Publication 1228
NASM3-20 technical data
NASM525 technical data
AN970-3 Washer technical data

Aerodynamic Load Basis

Fluid-Dynamic Drag - S. F. Horner

Composite Fabrication & Material Data

MIL-C-9084 - Style 7781 fiberglass fabric specification
EPOCAST 50-A1 / Hardener 9816 manufacturer technical data