NASA TN D-7400
Table I, printed p. 5 (PDF p. 10): NTO/Aerozine 50, O/F 1.6, 123 psia, 309.7 s, 5.08 kg/s, and 45.6 area ratio. The prose and table distinguish 3,500 lbf nominal/design thrust from a 3,450 lbf typical duty-cycle value.
NASA recordThe outputs, before you install anything
Below: the complete Apollo LMAE reference package with its comparison against NASA records, and screenshots of the report, DXF, injector, and flowfield outputs. Everything comes from one solved case with the assumptions and warnings visible.
The LMAE reference case
The Apollo Lunar Module Ascent Engine is a documented pressure-fed, fixed-thrust, hypergolic vacuum engine. This versioned reference case uses its intended vacuum condition and compares RPL Engine Workbench outputs with primary NASA and contractor records. It is reference-case agreement, not whole-solver validation.
| Quantity | Historical | RPL | Difference | Classification |
|---|---|---|---|---|
| Vacuum thrust | 15,568.8 N design target | 15,568 N | −0.8 N (−0.005%) | Entered target closure |
| Vacuum Isp | 309.7 s | 307.9659 s | −1.7341 s (−0.5599%) | Independent output |
| Total flow | 5.08 kg/s | 5.15477 kg/s | +0.07477 kg/s (+1.472%) | Qualified output |
| Effective c* | 1,724.7108 m/s injector test reduction | 1,616.64 m/s | −108.0708 m/s (−6.2660%) | Qualified diagnostic only |
Historical inputs: NTO/Aerozine 50, O/F 1.6, chamber pressure 0.848 MPa (123 psia), expansion ratio 45.6, target vacuum thrust 15,568 N, and ambient pressure 0 Pa. These are case inputs, not independent validation successes.
Interpretation: the −0.56% Isp difference is useful reference-case agreement, but it does not isolate model accuracy because several engine conditions were supplied and efficiency factors were assumed. The −6.27% c* comparison is less comparable: Boeing's figure is a production-injector acceptance-test reduction, not a flight-engine c*. The package explains the source-pressure, hardware, and reduction-convention differences rather than hiding the unfavorable result.
Four-file package: a five-page START_HERE comparison, the 11-page app-generated report, the matching millimeter DXF, and the portable input-only project. No sea-level case, synthetic calibration demo, logs, screenshots, duplicate outputs, or empty files.
SHA-256: a020047d5209d27156235a0d0903d2b42cedd35c09ccf55127fcf5d1237f5875 · 217,047 bytes · generated 2026-07-10 with RPL Engine Workbench 1.0.0. Preliminary engineering software output — not flight qualification or manufacturing authorization.
NASA TN D-7082
Printed pp. 1–2 (PDF pp. 5–6): the ascent system's 3,500 lbf requirement, O/F 1.6, and nitrogen tetroxide with equal hydrazine/UDMH fuel.
NASA recordBoeing D2-117060-2
Section 4.5.1, printed p. 70 (PDF p. 78), and Appendix B: production Rocketdyne injector c*. Section 2.3.2.3 documents why differing pressure-tap conventions can prevent direct comparison.
NASA recordSources retrieved 2026-07-10. Historical Apollo data are used as independent public references. NASA, Boeing, Rocketdyne, and the Apollo program do not sponsor or endorse RPL Engine Workbench.
Captured from the app
In trial mode you can reproduce the solve and export a watermarked PDF report; DXF and VTK export require a paid license.
Provenance
The reference project above was generated by the same signed, notarized build served on the download page; build identity and per-file SHA-256 checksums are inside the package. The validation page explains the bounded reference cases and automated checks behind the public examples. Results still require qualified review, appropriate test data, and your team's safety and compliance process.