Component Weight Estimation

Component Weight Estimation (vehicle-design/sizing/component-weight-estimation)

Use when the task is predicting the four airframe structural group masses at the class-II level from geometry and design loading: the wing group, the horizontal tail group, the vertical tail group and the fuselage group, each from a published closed-form statistical regression on its planform or body dimensions, sweep, thickness ratio, dynamic pressure, design load factor and design gross weight. This leaf implements the class-II statistical group-weight prediction method of the weight estimation chapter of Raymer, Aircraft Design: A Conceptual Approach, paraphrased in this leaf's own notation (summary-only, reference-only). It pairs with vehicle-design/sizing/wing-planform-sizing, tail-sizing and fuselage-sizing for the geometry inputs, vehicle-design/sizing/ fuel-tank-sizing for the in-wing fuel mass, and hands its group masses downstream to vehicle-design/sizing/weight-estimation and vehicle-design/mass-properties/mass-budget as given component inputs.

Domain quick reference

• Ultimate load factor: N_ult = 1.5 * nz_limit, the FAR-25.303 factor of safety between the design limit maneuvering load factor and the ultimate value used in every regression.
• Wing group (lb): W_w = 0.036 * S_w^0.758 * W_fw^0.0035 * (A_w / cos^2 L_w)^0.6 * q^0.006 * lam_w^0.04 * (100 * tc_w / cos L_w)^-0.3 * (N_ult * W0)^0.49.
• Horizontal tail group (lb): W_ht = 0.016 * (N_ult * W0)^0.414 * (A_ht / cos^2 L_ht)^0.168 * q^0.043 * S_ht^0.896 * (100 * tc_ht / cos L_ht)^-0.12 * lam_ht^0.02.
• Vertical tail group (lb): W_vt = 0.073 * (1 + 0.2 * H_t) * (N_ult * W0)^0.376 * q^0.122 * S_vt^0.873 * (100 * tc_vt / cos L_vt)^-0.49 * A_vt^0.357 * (lam_vt / cos^2 L_vt)^0.039, with H_t the T-tail location factor (0.0 fuselage-mounted, 1.0 T-tail).
• Fuselage group (lb): W_fus = 0.052 * S_f^1.086 * (N_ult * W0)^0.177 * (L_f / D_f)^-0.072 * q^0.241, plus the additive pressurization penalty W_press = 11.9 * (V_p * dp)^0.271 in lb when a pressurized volume and a pressure differential are both supplied.
• Every regressor is SI in, kg out; the regressions themselves run in the published lb / ft^2 / psf closed forms and convert at the boundary with fixed unit constants (KG_TO_LB, M_TO_FT, PA_TO_PSF, PA_TO_PSI), so every published coefficient stays visible.
• The four regressions are sublinear in (N_ult * W0), every exponent below 0.5, so the airframe group total scales slower than MTOW as the design grows.
• FAR-25 (14 CFR Part 25) and CS-25 set the certification context that the ultimate-over-limit factor of safety comes from.

Workflow

1. Collect the design loading: MTOW in kg, the design limit maneuvering load factor, and the cruise dynamic pressure at the design point in Pa.
2. Evaluate the wing group with wing_group_weight on the wing planform area, aspect ratio, taper ratio, thickness to chord, quarter-chord sweep and the in-wing fuel mass.
3. Evaluate the horizontal tail group with horizontal_tail_group_weight and the vertical tail group with vertical_tail_group_weight on the respective tail planforms, passing the T-tail location factor for a T-tail configuration.
4. Evaluate the fuselage group with fuselage_group_weight on the wetted area, length and equivalent diameter, supplying the pressurized volume and cabin pressure differential together when the fuselage is pressurized.
5. Sum the four group masses into the airframe group total with airframe_group_total.
6. Divide each group mass and the total by MTOW to get the per-group and total fractions that feed the weight statement, and hand them downstream to the weight-and-balance and mass-budget consumers.
7. Confirm the deterministic checks with the contract test scripts/test_component_weight_estimation.py.

Worked example

180-seat narrowbody transport, all inputs SI: MTOW 79000.0 kg, design limit maneuvering load factor 2.5 (N_ult 3.75), cruise dynamic pressure 12000.0 Pa.

• Wing: planform area 125.0 m^2, aspect ratio 10.0, taper ratio 0.25, thickness to chord 0.11, quarter-chord sweep 25.0 degrees, in-wing fuel 14000.0 kg. wing_group_weight gives 5835.319661 kg.
• Horizontal tail: area 32.0 m^2, aspect ratio 6.0, taper ratio 0.30, thickness to chord 0.09, sweep 30.0 degrees. horizontal_tail_group_weight gives 461.175701 kg.
• Vertical tail: area 26.0 m^2, aspect ratio 1.8, taper ratio 0.30, thickness to chord 0.12, sweep 35.0 degrees, fuselage-mounted (t_tail 0.0). vertical_tail_group_weight gives 438.062259 kg; the T-tail variant (t_tail 1.0) gives 525.674710 kg, exactly 1.2 times the fuselage-mounted value.
• Fuselage: wetted area 405.0 m^2, length 39.5 m, equivalent diameter 3.9 m, pressurized volume 300.0 m^3, cabin pressure differential 55158.0 Pa. fuselage_group_weight gives 7362.993775 kg; the unpressurized variant gives 7246.112306 kg, so the pressurization penalty is 116.881468 kg (about 1.6 percent of the fuselage group).
• airframe_group_total gives 14097.551395 kg. As fractions of the 79000.0 kg MTOW: wing 0.073865, horizontal tail 0.005838, vertical tail 0.005545, fuselage 0.093202, four-group total 0.178450, which sits well below the class-I transport band lower bound of 0.42 that the weight-estimation sibling checks.
• Doubling the ultimate load factor scales the wing by 2^0.49 = 1.404445, the horizontal tail by 2^0.414 = 1.332375, the vertical tail by 2^0.376 = 1.297739 and the unpressurized fuselage by 2^0.177 = 1.130531. Doubling the dynamic pressure scales the same four groups by 2^0.006, 2^0.043, 2^0.122 and 2^0.241.

Verification

• Confirm the worked-example values above to within 1e-6 relative for each group and the total.
• Confirm every group mass is positive, below the total, and the total is below MTOW; confirm the wing and fuselage fractions of MTOW fall in [0.05, 0.15], each empennage fraction in [0.004, 0.03], and the four-group fraction in [0.10, 0.30].
• Confirm the exact power-law identities: doubling the wing planform area, aspect ratio, taper ratio, in-wing fuel, tail areas, fuselage wetted area and fuselage length-to-diameter ratio scale each group by 2 raised to its published exponent, within 1e-9 relative.
• Confirm the T-tail factor is exactly 1.2 times the fuselage-mounted value within 1e-12 relative.
• Confirm the pressurization penalty is the difference between the pressurized and unpressurized fuselage mass, scales as 2^0.271 when the pressure differential doubles, and is unchanged when the load factor or dynamic pressure changes.
• Confirm ValueError rejection of a non-positive MTOW, limit load factor, dynamic pressure, planform area, aspect ratio, taper ratio, thickness to chord, sweep angle, wetted area, fuselage length and diameter, in-wing fuel mass, and negative t_tail or group mass; a thickness to chord at or above 1.0, a sweep at or above 90 degrees, and exactly one pressurization parameter supplied without the other.
• Run the contract test offline: python3 scripts/test_component_weight_estimation.py (deterministic, identical outputs under both interpreters).

Related leaves

• vehicle-design/sizing/wing-planform-sizing: the wing planform geometry that feeds the wing group regression.
• vehicle-design/sizing/tail-sizing: the tail planform geometry that feeds the horizontal and vertical tail group regressions.
• vehicle-design/sizing/fuselage-sizing: the fuselage dimensions that feed the fuselage group regression.
• vehicle-design/sizing/fuel-tank-sizing: the in-wing fuel mass input to the wing group regression.
• vehicle-design/sizing/weight-estimation: consumes the four group masses as given component weights for moments, CG and empty-weight fraction band checks.
• vehicle-design/mass-properties/mass-budget: rolls the group masses into the subsystem mass budget with growth allowance and margin.

Pitfalls

• Feeding the limit load factor instead of the ultimate load factor into a regression: every group uses N_ult = 1.5 * nz_limit, so passing the limit value understates every group mass.
• Calling wing_group_weight with no in-wing fuel: the regression requires a positive fuel_in_wing_kg; a fuel-less wing is outside its domain and raises ValueError rather than silently zeroing the fuel term.
• Supplying only one of the pressurization parameters: the fuselage regression requires the pressurized volume and the pressure differential together, or neither; one without the other raises ValueError.
• Reading a group mass as a takeoff weight, an empty weight or a total aircraft mass: the four groups are the airframe structural share only, always below the class-I empty-weight fraction band that the weight-estimation sibling checks against given component weights.
• Treating the omitted tail-arm decorrelation factor as missing precision: the fuselage regression here omits the source table's small tail-arm term (below class-II resolution), changing the fuselage mass by under 1 percent against the full published form.

Behavior contract (gate 3)

Run the deterministic contract test (stdlib unittest, offline):

python3 scripts/test_component_weight_estimation.py

The test exercises workflow steps 2 through 6: the wing, horizontal tail, vertical tail and fuselage group regressions at the worked example, the airframe group total and its MTOW fractions, the T-tail factor, the pressurization decomposition, the load-factor and dynamic pressure power laws, the geometric power laws, the sweep penalty, and ValueError rejection of every non-physical input enumerated above.

Compliance

• Standards referenced, not reproduced: FAR-25 (14 CFR Part 25) is US government work (public domain) and CS-25 is a free EASA download; the class-II statistical group-weight method above is standard engineering methodology, summary-only per standards-map.yaml.
• compliance: STANDARDS-REF, gated: false.